avionic emb 145-3

Business and Commuter Aviation SystemsHoneywell Inc.Box 29000Phoenix, Arizona 85038

22-05-14TITLE PAGE T-1

PRINTED IN U.S.A. PUB. NO. A15-1146-065 1 NOVEMBER 1996

PRIMUS® 1000 IntegratedAvionics System

Embraer 145

SystemMaintenance Manual

Volume III — System Test and Fault Isolation

SYSTEMMAINTENANCEMANUALEMBRAER 145

22-05-21TITLE PAGE T-2

Copyright 1996 Honeywell Inc.All Rights Reserved 1 SEPTEMBER 1996

PROPRIETARY NOTICE

This document and the information disclosed herein are proprietary data of Honeywell Inc.Neither this document nor the information contained herein shall be used, reproduced, ordisclosed to others without the written authorization of Honeywell Inc., except to the extentrequired for installation or maintenance of recipient’s equipment.

NOTICE - FREEDOM OF INFORMATION ACT (5 USC 552) ANDDISCLOSURE OF CONFIDENTIAL INFORMATION GENERALLY (18 USC 1905)

This document is being furnished in confidence by Honeywell Inc. The information disclosedherein falls within exemption (b) (4) of 5 USC 552 and the prohibitions of 18 USC 1905.

S96


Page RR-1Nov 1/1996

Use or disclosure of information on this page is subject to the restrictions on the title page of this document.

22-05-14

RECORD OF REVISIONS – VOLUME III

Upon receipt of a revision, insert the latest revised pages and dispose of superseded pages. Enterrevision number and date, insertion date, and the incorporator’s initials on the Record of Revisions.The typed initials HI are used when Honeywell Inc. is the incorporator.

RevisionNumber

RevisionDate

InsertionDate By

RevisionNumber

RevisionDate

InsertionDate By


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LIST OF EFFECTIVE PAGES - VOLUME III

Original … 0 … Nov 1/1996

SUBHEADING AND PAGE REVISION SUBHEADING AND PAGE REVISION

TitleT-1 0T-2 0

Record of RevisionsRR-1 0RR-2 0

List of Effective PagesLEP-1 0LEP-2 0

Section 7System Test and Fault Isolation

7-1 07-2 07-3 07-4 07-5 07-6 07-7 07-8 07-9 07-10 07-11 07-12 07-13 07-14 07-15 07-16 07-17 07-18 07-19 07-20 07-21 07-22 07-23 07-24 07-25 07-26 07-27 07-28 07-29 07-30 07-31 07-32 07-33 07-34 0

7-35 07-36 07-37 07-38 07-39 07-40 07-41 07-42 07-43 07-44 07-45 07-46 07-47 07-48 07-49 07-50 07-51 07-52 07-53 07-54 07-55 07-56 07-57 07-58 07-59 07-60 07-61 07-62 07-63 07-64 07-65 07-66 07-67 07-68 07-69 07-70 07-71 07-72 07-73 07-74 07-75 07-76 07-77 07-78 07-79 07-80 07-81 07-82 0


SUBHEADING AND PAGE REVISION SUBHEADING AND PAGE REVISION

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Section 7 (cont)System Test and Fault Isolation

7-83 07-84 07-85 07-86 07-87 07-88 07-89 07-90 07-91 07-92 07-93 07-94 07-95 07-96 07-97 07-98 07-99 07-100 07-101 07-102 07-103 07-104 07-105 07-106 07-107 07-108 07-109 07-110 07-111 07-112 07-113 07-114 07-115 07-116 07-117 07-118 07-119 07-120 07-121 07-122 07-123 07-124 0


-1Nov 1/1996


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SECTION 7

SYSTEM TEST AND FAULT ISOLATION

This section provides a Ground Maintenance Test (GMT) in Table 7-1 for checking the units of thePRIMUS® 1000 system for correct installation and proper operation. Table 7-1 also contains faultverification and corrective action procedures.

NOTE: The procedures provided in Table 7-1, Ground Maintenance Test (GMT) Procedures, arebased on Engineering Bulletin EB7023389, Rev -.


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Table 7-1. Ground Maintenance Test (GMT) Procedures

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TABLE OF CONTENTS

Para Title Page

SECTION I IC-600 SELF TEST ............................................................................................... 7-9

1. INTRODUCTION ............................................................................................................................ 7-9

1.1 SCOPE..............................................................................................................................................7-9

1.2 REFERENCE DOCUMENTS ....................................................................................................................7-9

1.3 TERMS AND DEFINITIONS .....................................................................................................................7-9

2. SYSTEM BENCH REQUIREMENTS ................................................................................................. 7-10

3. TEST MODES............................................................................................................................. 7-10

3.1 IN-AIR-INITIATED TEST...................................................................................................................... 7-10

3.2 ON-GROUND INITIATED TEST ............................................................................................................. 7-123.2.1 FAILURE MODE ANNUNCIATION AND FAMILIARIZATION .................................................................. 7-12

3.2.1.1 EICAS TEST............................................................................................................. 7-163.2.2 INTERNAL BUILT-IN TEST ......................................................................................................... 7-18

3.3 MAINTENANCE TEST MODE ................................................................................................................ 7-203.3.1 INDEX PAGES ......................................................................................................................... 7-23

3.3.1.1 INDEX PAGE 1 OF 4 (RA SETTING 600) ........................................................................ 7-233.3.1.2 INDEX PAGE 2 OF 4 (RA SETTING 610) ........................................................................ 7-243.3.1.3 INDEX PAGE 3 OF 4 (RA SETTING 620) ........................................................................ 7-253.3.1.4 INDEX PAGE 4 OF 4 (RA SETTING 630) ........................................................................ 7-26

3.3.2 HARDWARE PAGES ................................................................................................................. 7-273.3.2.1 HW ID 1 (RA SETTING 640)....................................................................................... 7-273.3.2.2 HW 2 (RA SETTING 650) ........................................................................................... 7-28

3.3.3 DISPLAY UNIT PAGES .............................................................................................................. 7-293.3.3.1 DISPLAY UNITS - PILOT’S SIDE (RA SETTING 660) ......................................................... 7-293.3.3.2 DISPLAY UNITS - COPILOT’S SIDE (RA SETTING 660C) .................................................. 7-30

3.3.4 EVENT CODES (RA SETTING 670) ............................................................................................ 7-313.3.5 POWER AND INTERNAL TEMPERATURE PAGE (RA SETTING 680) ................................................... 7-333.3.6 AHRS DATA (RA SETTING 690) .............................................................................................. 7-343.3.7 DC VARIABLES (RA SETTING 700) ........................................................................................... 7-353.3.8 MICRO A IR DATA COMPUTER (RA SETTING 710) ........................................................................ 7-363.3.9 DC-550 PORT (RA SETTING 720) ............................................................................................ 7-383.3.10 WEATHER RADAR (RA SETTING 730)...................................................................................... 7-413.3.11 DISCRETES 1 (RA SETTING 740) ............................................................................................ 7-423.3.12 DISCRETES 2 (RA SETTING 750) ............................................................................................ 7-433.3.13 DISCRETES 3 (RA SETTING 760) ............................................................................................ 7-443.3.14 DISCRETES 4 (RA SETTING 770) ............................................................................................ 7-453.3.15 DISCRETES 5 (RA SETTING 780) ............................................................................................ 7-463.3.16 DISCRETES 6 (RA SETTING 790) ............................................................................................ 7-473.3.17 DISCRETES 7 (RA SETTING 800) ............................................................................................ 7-483.3.18 DISCRETES 8 (RA SETTING 810) ............................................................................................ 7-493.3.19 DISCRETES 9 (RA SETTING 820) ............................................................................................ 7-503.3.20 LAMP TEST 9 (RA SETTING 830)............................................................................................ 7-51



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3.3.21 ARINC 1 (RA SETTING 840) ................................................................................................. 7-533.3.22 ARINC 2 (RA SETTING 850) ................................................................................................. 7-543.3.23 ARINC 3 (RA SETTING 860) ................................................................................................. 7-553.3.24 ARINC 4 (RA SETTING 870) ................................................................................................. 7-563.3.25 RADIO SYSTEM BUS (RA SETTING 880) .................................................................................. 7-573.3.26 CHECKLIST LOADING (RA SETTING 890) ................................................................................. 7-583.3.27 SYSTEM CONFIGURATION (RA SETTING 900)............................................................................ 7-593.3.28 BLANK PAGE (RA SETTING 910) ............................................................................................ 7-59

4. EVENT CODES ........................................................................................................................... 7-60

SECTION II STANDARD GROUND TEST PROCEDURE .............................................................. 7-73

1. INTRODUCTION .......................................................................................................................... 7-73

2. DISCRETE INPUT VERIFICATIONS ................................................................................................. 7-73

3. DU-870 (DISPLAY UNIT) ............................................................................................................ 7-74

4. DC-550 DISPLAY CONTROLLER.................................................................................................. 7-74

5. IC-600 INTEGRATED COMPUTER ................................................................................................. 7-75

6. AZ-850 MICRO AIR DATA COMPUTER ......................................................................................... 7-75

7. AH-800 ATTITUDE AND HEADING REFERENCE UNIT...................................................................... 7-75

8. PRIMUS II RADIOS ................................................................................................................... 7-75

9. AA-300 RADIO ALTIMETER......................................................................................................... 7-76

10. DAU DATA ACQUISITION UNIT .................................................................................................. 7-76

11. TCAS TRAFFIC COLLISION AVOIDANCE SYSTEM ........................................................................ 7-76

11.1 PROCEDURES ................................................................................................................................ 7-76

12. PRIMUS 650/660 WEATHER RADAR ........................................................................................ 7-78

12.1 PROCEDURES ................................................................................................................................ 7-78

13. GC-550 FLIGHT GUIDANCE CONTROLLER.................................................................................. 7-80

14. PC-400 AUTOPILOT CONTROLLER ............................................................................................ 7-80

15. SM-200.................................................................................................................................. 7-80

16. AUTOPILOT AND ELECTRIC TRIM ............................................................................................... 7-81



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SECTION III EXPANDED GROUND TEST PROCEDURE ............................................................. 7-83

1. INTRODUCTION .......................................................................................................................... 7-83

2. PRIMUS 650/660 WEATHER RADAR SYSTEM CHECK .................................................................. 7-83

2.1 PRELIMINARY CONTROL SETTINGS...................................................................................................... 7-83

2.2 PRECAUTIONS.................................................................................................................................. 7-83

2.3 TEST PATTERN............................................................................................................................... 7-83

2.4 STABILIZATION INPUT........................................................................................................................ 7-842.4.1 PRELIMINARY CHECKS............................................................................................................. 7-842.4.2 ELEVATION FEEDBACK CHECK/ADJUSTMENT .............................................................................. 7-842.4.3 PITCH GAIN ADJUSTMENT/CHECK ............................................................................................. 7-842.4.4 ROLL GAIN ADJUSTMENT/CHECK .............................................................................................. 7-84

2.5 FAULT MONITORS............................................................................................................................. 7-85

3. EFIS CHECK ............................................................................................................................. 7-85

3.1 PRELIMINARY CONTROL SETTINGS...................................................................................................... 7-85

3.2 DU-870 DISPLAY UNITS ................................................................................................................... 7-85

3.3 DC-550 DISPLAY CONTROLLER ......................................................................................................... 7-86

3.4 DIMMING CONTROLLERS.................................................................................................................... 7-87

3.5 REVERSIONARY PANEL ..................................................................................................................... 7-87

3.6 MFD BEZEL BUTTONS AND ROTARY SET KNOB ................................................................................... 7-87

4. MADC CHECK .......................................................................................................................... 7-89

4.1 NORMAL MODE ................................................................................................................................ 7-89

4.2 SELF-TEST MODE ............................................................................................................................ 7-90

5. RADIO ALTIMETER TEST ............................................................................................................. 7-91

6. AHRS CHECK ........................................................................................................................... 7-91

6.1 AHRS POWER-UP CHECKOUT........................................................................................................... 7-92

6.2 AHRS TEST MODE .......................................................................................................................... 7-92

6.3 AHRS ATTITUDE OUTPUTS CHECK..................................................................................................... 7-92

6.4 AHRS FLUX VALVE CHECK............................................................................................................... 7-92

6.5 VERIFICATION OF MEMORY MODULE PROGRAMMING.............................................................................. 7-93

7. PRIMUS II RADIOS CHECK ........................................................................................................ 7-95

7.1 BUS OPERATION CHECKOUT .............................................................................................................. 7-95

7.2 UNIT POWER ON SELF TEST (POST).................................................................................................. 7-97

7.3 AIRCRAFT INTERFACE/RAMP TESTS .................................................................................................... 7-987.3.1 COM 1 TESTS ....................................................................................................................... 7-987.3.2 TRANSPONDER 1 TESTS .......................................................................................................... 7-997.3.3 NAV 1 TESTS ...................................................................................................................... 7-100

7.3.3.1 VOR TESTS ............................................................................................................ 7-1007.3.3.2 LOC 1 TESTS ......................................................................................................... 7-1017.3.3.3 GLIDE SLOPE 1 TESTS ............................................................................................. 7-1017.3.3.4 MARKER BEACON 1 TESTS ....................................................................................... 7-101

7.3.4 DME 1 TESTS...................................................................................................................... 7-1027.3.5 ADF 1 TESTS ...................................................................................................................... 7-102



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7.3.6 RMU TESTS ........................................................................................................................ 7-1037.3.6.1 DIMMING CHECK ...................................................................................................... 7-1037.3.6.2 POWER ON/OFF CHECK .......................................................................................... 7-103

7.3.7 TUNING BACKUP CONTROL HEAD (CLEARANCE DELIVERY HEAD - CDH) TESTS............................ 7-104

7.4 SYSTEM STATUS MESSAGE CHECKS ................................................................................................. 7-104

7.5 IN-FLIGHT TESTS ........................................................................................................................... 7-1057.5.1 AUDIO PANEL FLIGHT TEST PROCEDURES ................................................................................ 7-105

8. TRAFFIC COLLISION AVOIDANCE SYSTEM (TCAS) ...................................................................... 7-106

8.1 TCAS CHECK ............................................................................................................................... 7-106

9. EICAS CHECK ........................................................................................................................ 7-111

9.1 REVERSIONARY CONTROL ............................................................................................................... 7-111

9.2 CREW ALERTING SYSTEM ................................................................................................................ 7-1129.2.1 TONE GENERATOR AND MESSAGE ACKNOWLEDGMENT............................................................... 7-1129.2.2 TAKE OFF INHIBIT FUNCTION................................................................................................... 7-1139.2.3 RMU ENGINE BACKUP .......................................................................................................... 7-1139.2.4 RMU NAVIGATION BACKUP.................................................................................................... 7-113

10. FLIGHT DIRECTOR TEST ......................................................................................................... 7-114

11. AUTOPILOT TEST ................................................................................................................... 7-122



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List of Figures(Section I)

Figure 3.1.1. DC-550 Display Controller .................................................................................. 7-10Figure 3.1.2. PFD In-Air Initiated Test Mode ........................................................................... 7-11Figure 3.2.1. PFD On-Ground Initiated Test Mode................................................................... 7-13Figure 3.2.2. MFD On-Ground Intitiated Test Mode ................................................................. 7-15Figure 3.2.3. EICAS On-Ground Initiated Test Mode ............................................................... 7-17Figure 3.2.4. Interface Self-Test Display ................................................................................. 7-18Figure 3.3.1. Maintenance Display Formats on the DU-870...................................................... 7-22Figure 3.3.2. Index Page 1 of 4 .............................................................................................. 7-23Figure 3.3.3. Index Page 2 of 4 .............................................................................................. 7-24Figure 3.3.4. Index Page 3 of 4 .............................................................................................. 7-25Figure 3.3.5. Index Page 4 of 4 .............................................................................................. 7-26Figure 3.3.6. Hardware Identification Page 1........................................................................... 7-27Figure 3.3.7. Hardware Identification Page 2........................................................................... 7-28Figure 3.3.8. Display Unit Page (Pilot’s Side) .......................................................................... 7-29Figure 3.3.9. Display Unit Page (Copilot’s Side) ...................................................................... 7-30Figure 3.3.10. Event Code Page ............................................................................................. 7-31Figure 3.3.11. Power/Internal Temperature Page .................................................................... 7-33Figure 3.3.12. AHRS Data Page ............................................................................................. 7-34Figure 3.3.13. DC Variables Page........................................................................................... 7-35Figure 3.3.14. Air Data Page .................................................................................................. 7-36Figure 3.3.15. DC-550 Port Page............................................................................................ 7-38Figure 3.3.16. Weather Radar Page ....................................................................................... 7-41Figure 3.3.17. Discretes 1 ...................................................................................................... 7-42Figure 3.3.18. Discretes 2 ...................................................................................................... 7-43Figure 3.3.19. Discretes 3 ...................................................................................................... 7-44Figure 3.3.20. Discretes 4 ...................................................................................................... 7-45Figure 3.3.21. Discretes 5 ...................................................................................................... 7-46Figure 3.3.22. Discretes 6 ...................................................................................................... 7-47Figure 3.3.23. Discretes 7 ...................................................................................................... 7-48Figure 3.3.24. Discretes 8 ...................................................................................................... 7-49Figure 3.3.25. Discretes 9 ...................................................................................................... 7-50Figure 3.3.26. Lamp Test ....................................................................................................... 7-51Figure 3.3.27. ARINC #1 ........................................................................................................ 7-53Figure 3.3.28. ARINC #2 ........................................................................................................ 7-54Figure 3.3.29. ARINC #3 ........................................................................................................ 7-55Figure 3.3.30. ARINC #4 ........................................................................................................ 7-56Figure 3.3.31. Radio System Bus............................................................................................ 7-57Figure 3.3.32. Check Loading Page ........................................................................................ 7-58Figure 3.3.33. System Configuration ....................................................................................... 7-59Figure 3.3.34. Blank Page ...................................................................................................... 7-59



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List of Figures (cont)(Section II)

Figure 11.1.1. TCAS Test Display........................................................................................... 7-77Figure 12.1.1. PRIMUS® 650 MFD Weather Radar Test Mode ................................................ 7-78Figure 12.1.2. PRIMUS® 660 MFD Weather Radar Test Mode ................................................ 7-79

(Section III)

Figure 3.6.1. MFD Bezel Button Menu Tree............................................................................. 7-88Figure 8.1.1. Antenna Test Points (Rack Side) ...................................................................... 7-107

List of TablesTable 4.1. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Power Up Event

Codes That Prevent Autopilot Engagement)....................................................... 7-60Table 4.2. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Continuous Event

Codes That Disengage the AP and YD Autopilot Engagement) ........................... 7-66Table 4.3. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Power Up Event

Codes) ............................................................................................................. 7-68Table 4.4. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes (Event Codes That

Record An Action) ............................................................................................ 7-70Table 4.5. Autopilot and Interloop Monitors Event Codes (Power Up Event Codes That Prevent

Autopilot Engagement)...................................................................................... 7-71Table 4.6. Autopilot and Interloop Monitors Event Codes (Continuous Event Codes That Prevent

Autopilot Engagement)...................................................................................... 7-72Table 4.7. Autopilot and Interloop Monitors Event Codes Execution Failure.............................. 7-72Table 8.1. CU Fault Reporting and Corrective Actions................................................................... 7-108Table 8.2. Antenna Wiring Resistance....................................................................................... 7-109Table 8.3. CU Self-Test Execution............................................................................................ 7-110Table 10.1. Flight Director Test ............................................................................................ 7-114Table 11.1. Autopilot Test .................................................................................................... 7-122



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SECTION IIC-600 SELF TEST

1. INTRODUCTION

1.1 SCOPE

This specification outlines the self-test capabilities of the IC-600. It decribes in detail how toexecute each test from an aircraft and/or system bench. The available self-tests within theIC-600 include the In-Air-Initiated Test (Weight Off Wheels) and the On-Ground-Initiated Test(Weight On Wheels).

1.2 REFERENCE DOCUMENTS

Embraer, EMB 145 System and Software Requirements Document, PS7021214

1.3 TERMS AND DEFINITIONS

TERM DEFINITION

AC Aircraft

ADC Air Data Computer

ADI Attitude Direction Indicator

BIT Built In Test

DU Display Unit

HSI Horizontal Situation Indicator

ICB Integrated Computer Bus

LRN Long Range Nav

MFD Multi-Function Display

RSB Radio System Bus

RX Receiver

TX Transmitter

WOW Weight On Wheels

WSP Word Sequence Position

WX Weather Radar

LRU Line Replaceable Unit



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2. SYSTEM BENCH REQUIREMENTS

In order to perform the IC-600 test on the system bench, the following requirements must be met:

• The 115 V ac 400 Hz bench power is ON.

• The 28 V dc bench power is ON.

• Factory test identification discretes are all open. (These discretes into the IC-600 should not bewired in the aircraft or system bench. They are intended for use in a factory repair environmentonly.)

• The IC-600, displays, and remaining avionics must be powered on to optimize self test results.

3. TEST MODES

3.1 IN-AIR-INITIATED TEST

An in flight (Weight Off Wheels) test may be initiated by pushing and holding the TEST button onthe DC-550 (Figure 3.1.1).

NOTE: This test is inhibited when Glide Slope is captured.

The test results in the active FD mode annunciations on the PFD being replaced by a magenta“TEST” annunciation (Figure 3.1.2). The Radio Altimeter will also be commanded into the testmode resulting in the display of a fixed radio altitude of 100 ft. The MFD and EICAS will notchange during this test.

No other visible testing of the IC-600 will be executed during in this “in-air” test. However,background testing of the IC-600 is continually in process verifying such parameters as powersupply levels, analog to digital functions, software/hardware validity, and other parameters.

BRG BRG

NAV FMSET

OFF

NAV 1

ADF

GSPDTTG

FULLWX

FMS OFF

NAV 2

ADF

FMS

RA TEST

AD-50629-R1@

Figure 3.1.1. DC-550 Display Controller



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NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS

M

MIN1

260

25

260

24240

220

200

280

TEST

AP YD

10 10

20 20

10 10

20 20

145 00

14500

14320

8000

1

9 R2

FMS

14000

MAX

SPD

AOA

ATT2

AD-51487@

001HDG

359

E

W

21

12

30

15

33

S

24

.410 M

VOR1

ADF2

200 RA

3

1

2

3000

0

1

2

3

25

29.92 IN

N3

6

CRS

TGTTTG5MIN

13.1 NM

350

MAG2

Figure 3.1.2. PFD In-Air Initiated Test Mode



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3.2 ON-GROUND INITIATED TEST

The IC-600 On-Ground (Weight On Wheels) initiated test has two modes of operation. Thesemodes are:

• Failure Mode Annunciation and Familiarization

• Initiated Test Mode

3.2.1 FAILURE MODE ANNUNCIATION AND FAMILIARIZATION

To initiate the test, push and hold the TEST button on the DC-550 (Figure 3.1.1). Failuremode annunciation and familiarization will occur for the first 5 to 6 seconds.

NOTE: This test is only functional on the ground. The radio altimeter test is functionalat all times except during GS CAP.

A magenta “TEST” will be displayed in the lateral arm location (top left center) as shown inFigure 3.2.1. The PFD test page format displays the following quantities as invalid:

• Pitch and Roll Attitude

• Vertical Deviation (Pointer and Scale)

• Barometric Correction

• Vertical Speed Set Digital Readout

• Altitude Preselect Digital Indicator

• Heading (Scale and digital readout)

• Lateral deviation (Pointer and Scale)

• Distance Digital Readout

• GSPD or TTG or Elapsed Time

• Selected course or desired track (pointer and digital readout)

• Selected heading (digital)

• Airspeed Indicator

• Mach Digital Readout

• Airspeed Set Digital Readout

• Altitude Indicator (Tape and Digits)

• Vertical Speed Indicator



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NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS

AP YD

VOR1

VOR2

ILS 1HDGFAIL

I M

- - -HDG

- - -

E

W

21

12

30

15

33

S

24

.- - -

1610

HDG

- - - - -

3

VS

1

2

0

1

2

3

TEST

O

N3

6

CRS

TTG- - -MIN

- - -

ATTFAIL

IAS

ALT

800 RA

TCASFAIL

RAILSCAS MSG

AD-51488@

ATT

WDSHEAR

AOA

M - -.- - IN

- - . - NM

Figure 3.2.1. PFD On-Ground Initiated Test Mode



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The following will be removed from the test page display:

• All bugs (airspeed, altitude, heading, drift, vertical speed)

• Flight Director information (command bars/cue, modes, FPA symbology)

• All pointers (bearing, To/From)

• Low Speed Awareness

• Vspeed bugs and digital readouts

• Vmo/Mmo

• Airspeed trend vector

• Altitude trend vector

The following display items will be forced on:

• All comparison monitors

• All three marker beacons

• Windshear warning annunciation

If heading is valid upon entering this test mode, the Heading source annunciation shall remaina valid annunciation - (DG1or DG2) or (MAG1 or MAG2) - and will NOT be forced to (HDG1orHDG2) as it normally would be when heading is invalid.

MFD Test Mode will always be displayed in MAP mode (Figure 3.2.2). The MFD test pageformat displays the following quantities as invalid:

• Heading

• Weather Radar Tilt

• True Airspeed

• Static Air Temp

• Total Air Temperature

• Nav Source distance to “TO” Waypoint

• Time to “TO” Waypoint

• Bezel Menu



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- - - -

N

S

33

15

3012

WE

246

21

3

- - - TAT

- - - SAT

- - - TAS

FMS

- -. -

- - -

NM

MIN

50 50

WX

AD-51489@

MENU INOP

DOORS

T/O MODE: - - - - - - - -

ENGINE

REF TO TEMP: - - -

REF A-ICE: - - -

OIL LVL- - - QT

C

HDGFAIL

- - - QT

Figure 3.2.2. MFD On-Ground Intitiated Test Mode



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The following symbology will be removed from the display:

• Heading Source Annunciation

• TCAS, Checklist or System Pages

• Weather Patch/Boundary Line

• Drift Bug

• Wind Vector

• Heading Select Bug

• Flight Plan Data

• Airports

• Navaids

• Designator Information

3.2.1.1 EICAS Test

A magenta “TEST” will be displayed on the PFD to signify that the On-Ground Test hasbeen initiated (Figure 3.2.1). The EICAS test page format displays the following quantitiesas invalid as shown in Figure 3.2.3:

• Landing Gear Status

• N1

• ITT

• N2

• Fuel Flow and Quantity

• Oil Pressure and Temperature.

• Vibration for Low and High Pressure Rotors

• Flaps

• Spoilers

• All Cabin and APU Parameters

• All trim values

The CAS message field will be filled with an X as shown in Figure 3.2.3.

The following symbology will be removed from the display:

• Reversion Annunciation

• Ignition Annunciation

• FADEC in Control Annunciation

• All Engine Bugs

• All Trim Bugs



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AD-51490@

LDG GEAR

FLAPS

SPLRS CLD

APU

CAB ALT

CAB P

CAB RATE

C

FT

PSI

FPM

%

VIB

LP HPTEMPPRESS

N1

ITT

N2 %%

FF PPHPPH

FQ LBLB

ROLL

YAW

PITCH

IGN IGN

OIL

- - - - - - - - - - - -

A

A

A A

ATTCS

- - - . -- - - . -

- - - - - - - -

- - . -

- - - -

- - - -

- - - . -- - - -

- - - -

- - - - - - - - - - - -

- - - - -

- - - - -

- - -- - -

- -

Figure 3.2.3. EICAS On-Ground Initiated Test Mode



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3.2.2 INTERNAL BUILT-IN TEST

The internal built-in test (IBIT) executes 5 - 7 seconds after the Failure Mode Annunciationand Familiarization if the TEST button remains pushed. The test exercises the IC-600sinternal interfaces using internal wraparounds. The test remains active as long as the TESTbutton is pushed and the PFD display format will cycle (every 10 seconds) between the IBIT1and IBIT2 test results. A short description of the hardware tested, followed by the statusannunciation, will be displayed as shown in Figure 3.2.4. Releasing the DC-550 test buttonterminates the test and resumes normal operation.

W A I T

P A S S

P A S S

P A S S

P A S S

P A S S

P A S S

P A S S

P A S S

P A S S

P A S S

I C - 6 0 0 I B I T

A U T O P I L O T

F L I G H T D I R E C T O R

S Y M B O L G E N E R A T O R

D I S P L A Y U N I T

A T T / H D G I N T E R F A C E

R A D A L T I N T E R F A C E

R A D I O I N T E R F A C E

D I S P L A Y C O N T R O L L E R

L A M P D R I V E R S

A D C D A T A

AD-51492@

Figure 3.2.4. Interface Self-Test Display

As each segment of hardware is tested the status is annunciated:

ANNUNCIATION COLOR DESCRIPTION

“VALD” Green Valid Data

“INVD” Red Invalid data is detected

“PASS” Green Successful end-around hardware test

“FAIL” Red Unsuccessful end-around hardware test

“N/A” White Not applicable/Not installed. Determined through theconfiguration discretes

A PASS/FAIL status indicates that the test is performed utilizing the manipulation of“end-around” hardware signals. These tests multiply the internal signals into the appropriatehardware. In the case of analog signals, the internal test voltages are used. Each test is runagain after the first failure to avoid erroneous FAIL determination.

A VALD/INVD (Valid/Invalid) status implies that the test is performed on incoming signals andNOT on internally wrapped signals.



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As each segment of hardware is tested the red FAIL/INVD annunciation will be replaced by agreen PASS or valid (VALD) if the hardware test is successful. It may be necessary to runthrough the tests more than once to isolate intermittents.

IBIT determines the status of the following items within the IC-600 as directed by theconfiguration:

• Autopilot status (if installed)

• Flight Director status

• Symbol Generator status

IBIT also tests the external or “blind sensors” are checked for valid data (a “blind sensor” isone whose failure is NOT apparent):

• Display Unit (DU-870)

• ATT/HDG interface

• Radio Alt Interface

• Radio Interface

• Display Controller

• Lamp Drivers

• ADC data (MADC).

The results of the Initiated Test are displayed on the PFD without further operator action.When the FAIL/INVD annunciation remains after performing the tests, the following actionsshould be accomplished:

• Verify the appropriate LRU providing the tested interface is installed and power is applied.

• Verify the wiring interface between the tested IC and LRU is correct.

Replace the tested LRU with the cross-side unit to verify if condition remains. If not, replacesuspect LRU and document the failed condition. Details of specific actions may be found inthe manufacturers aircraft fault isolation manual.



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3.3 MAINTENANCE TEST MODE

The maintenance displays provide the capability for the end user to display maintenance pageson the PFD. To enter the Maintenance Test Mode:

• Use the decision height (RA) knob on the DC-550 (Figure 3.1.1) to set a RA value between600 and 990 feet. Please note that four (4) INDEX pages at RA settings of 600, 610, 620,and 630 are provided as a reference to the available maintenance pages.

• Select decesion height knob on DC-550 for 600 or above on PFD.

• Push and hold TEST button/knob on the DC-550 for approximately 5 to 7 seconds .

• While holding TEST button, push ET button on the DC-550.

• Release the TEST button.

• The IC-600 should now be in the maintenance test mode and will remain in test (via software)until it is cancelled by another press of the TEST button OR selecting a RA setting below600.

In Maintenance Test mode, the MFD will be exactly the same as in the On-Ground-Initiated Testmode. The PFD will be replaced with the results of the various maintenance tests. Theseresults are displayed on several pages which are selected by the RA set knob on the DC-550.In addition to test results, there are also pages of sensor status and other miscellaneousinformation.

There will be a distinct page (and possible subpages) for each setting of RA, beginning with600 ft. and then for every 10 ft. above (i.e., 610, 620, etc). There are two sets of maintenancepages which are selectable based on the pushbutton selections. One set of maintenancepages, customer maintenance pages, are displayed for use in determining the valid operation ofthe IC-600. These pages are activated by pressing the TEST pushbutton approximately 5 to 7seconds and then pressing the ET pushbutton. The second set of maintenance pages,engineering maintenance pages, are displayed to assist in troubleshooting the IC-600 internaloperation. These engineering pages are activated by setting RA to 990 or greater and pressingthe FULL/WX pushbutton while in the customer maintenance pages. Once the engineeringpages have been entered, to display the customer mainenance pages upon re-entering theMaintenance Test Mode the ET pushbutton must be selected again or the engineering pages willbe displayed.



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NOTE: A description of the engineering maintenance pages are not provided in thisdocument. If you observe maintenance pages not described in the following textyou may have inadvertently entered engineering maintenance pages. These areuseful to engineering and testing personnel only.

The following maintenance pages are displayed based on the RA set value:

RA SETTING PAGE NAME

600 Index610 Index620 Index630 Index640 Hardware 1 Identification650 Hardware 2 Identification660 Pilot’s Display Unit Page660C Copilot’s Display Unit Page670 Event Codes680 Power/Internal Temperature690 AHRS Data700 DC Variables710 Air Data720 DC-550 Port730 Weather Radar740 Discretes 1750 Discretes 2760 Discretes 3770 Discretes 4780 Discretes 5790 Discretes 6800 Discretes 7810 Discretes 8820 Discretes 9830 Lamps Test840 ARINC 1850 ARINC 2860 ARINC 3870 ARINC 4880 Radio Sys Bus890 Checklist Loading900 Sys Config910 Blank Page



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The approximate location and size of Maintenance Display Formats on the DU-870 is shown inFigure 3.3.1. Each page is identified by a title (in green) on the first line. A textual description ofthe test result or data identification is displayed in white with the actual test results or data incyan. Operator instructions are displayed in white with key words displayed in amber.

AD-51493@

180 RA

HELPING YOU TO CONTROL YOUR WORLD

MAINTENANCEPAGEAREA

Figure 3.3.1. Maintenance Display Formats on the DU-870



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3.3.1 INDEX PAGES

Four Index pages presenting the Customer Index pages are provided.

3.3.1.1 Index Page 1 of 4 (RA Setting 600)

R A S E T

6 0 0

6 1 0

6 2 0

6 3 0

6 4 0

6 5 0

6 6 0

6 7 0

6 8 0

6 9 0

I N D E X P A G E S 1 O F 4

I N D E X

I N D E X

I N D E X

I N D E X

H W I D 1

H W I D 2

D I S P L A Y U N I T S

E V E N T C O D E S

P O W E R / I N T E R N A L T E M P

A H R S D A T A

- N E X T -

600 RAAD-51494@

Figure 3.3.2. Index Page 1 of 4



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R A S E T

7 0 0

7 1 0

7 2 0

7 3 0

7 4 0

7 5 0

7 6 0

7 7 0

7 8 0

7 9 0


D C V A R I A B L E S

A I R D A T A

D C - 5 5 0 P O R T

W E A T H E R R A D A R

D I S C R E T E S 1

D I S T R E T E S 2

D I S C R E T E S 3

D I S C R E T E S 4

D I S C R E T E S 5

D I S C R E T E S 6

- N E X T -

610 RAAD-51495@




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R A S E T

8 0 0

8 1 0

8 2 0

8 3 0

8 4 0

8 5 0

8 6 0

8 7 0

8 8 0

8 9 0


D I S C R E T E S 7

D I S C R E T E S 8

D I S C R E T E S 9

L A M P T E S T

A R I N C 1

A R N I C 2

A R N I C 3

A R N I C 4

R A D I O S Y S B U S

C H E C K L I S T L O A D I N G

- N E X T -

620 RAAD-51496@




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R A S E T

9 0 0

9 1 0

9 2 0

9 3 0

9 4 0

9 5 0

9 6 0

9 7 0

9 8 0

9 9 0


S Y S C O N F I G

B L A N K

B L A N K

B L A N K

B L A N K

B L A N K

B L A N K

B L A N K

B L A N K

B L A N K

630 RAAD-51497@




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3.3.2 HARDWARE PAGES

The Hardware Identification HW ID 1 page displays the following information:

• Base part number of IC-600 installed.

• Dash number of IC-600 installed.

• Serial number of IC-600 installed.

• Software mod status of the IC-600.

3.3.2.1 HW ID 1 (RA Setting 640)

7 0 1 7 0 0 0

- 5 0 1 0 X

X X X X X X X X

X X

H W I D 1

B A S E P A R T N O .

D A S H N O .

S E R I A L N O .

S W M O D S T A T U S

640 RAAD-51498@

Figure 3.3.6. Hardware Identification Page 1



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3.3.2.2 HW 2 (RA Setting 650)

A B C D E F G H

J K L M N P Q R S

T U V W X Y Z

A B C D E F G H

J K L M N P Q R S

T U V W X Y Z

A B C D E F G H

J K L M N P Q R S

T U V W X Y Z

H W 2

H W M O D S T A T U S

H W M O D S T A T U S A _

H W M O D S T A T U S B _

650 RAAD-51499@

Figure 3.3.7. Hardware Identification Page 2



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3.3.3 DISPLAY UNIT PAGES

This page displays the status of the Display Units. Data is wrapped back from display units,the MFD and EICAS on the HDLC bus and the PFD on a ARINC 429 channel. If any of thewrapped parameters do not agree, the DU WRAP will be set to FAIL. When this occurs, thefollowing acctions should be taken:

• Verify the appropriate DU is installed with power applied to the unit.

• Verify the appropriate Display Wrap Around interface between the IC-600 and thesuspect DU.

• Replace the suspect DU with a DU from another position and if the FAIL conditioncontinues to be observed, document the condition and remove the unit for repair.

PFD, MFD, and EICAS VALIDs remain true if the labels being sent from the DU to IC-600 arefresh. Each PFD (DUs #1 and #5) provide data via Arinc 429, while each MFD and theEICAS (DUs #2, 3, and 4) provide data via HDLC.

The DU Cooling Fan status is monitored by the Dispaly Units and the status is sent to theIC-600 via the Wrap Arounds. If the FAIL annunciation is displayed, the following actionsshould be taken:

• Verify that nothing is obstructing the fan inlet or outlet for the appropriate DU.

• Replace the suspect DU with a DU from another position and if the FAIL conditioncontinues to be observed, document the condition and remove the unit for repair.

3.3.3.1 Display Units - Pilot’s Side (RA Setting 660)

S T A T U S

P A S S

P A S S

P A S S

N / A

N / A

N / A

P A S S

P A S S

P A S S

D I S P L A Y U N I T S ( P )

D U # 1 W R A P

D U # 1 V A L I D

D U # 1 C O O L I N G F A N S

D U # 2 W R A P

D U # 2 V A L I D


D U # 3 W R A P

D U # 3 V A L I D


660 RAAD-51500@

Figure 3.3.8. Display Unit Page (Pilot’s Side)



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3.3.3.2 Display Units - Copilot’s Side (RA Setting 660C)

TO BE SUPPLIED

660C RAAD-51501@

Figure 3.3.9. Display Unit Page (Copilot’s Side)

LABEL DISPLAYED AS ANNUNCIATION PARAMETER LABEL

“DU #1 or 5 WRAP” “PASS” or “FAIL” PFD DU Wrap-Around See Note 1“DU #2,3,or 4 WRAP” “PASS” or “FAIL” MFD/EICAS DU Wrap-Around 356, Bits 16-30“DU#1 or 5 VALID” “PASS” or “FAIL” DU Data on HDLC See Note 2

DU#1 or 5 COOLING FANS” “PASS”or “FAIL1” Fan/airflow 0 test failure 350, Bit 22or “FAIL2” Fan/airflow 1 test failure 350, Bit 21or “FAIL 1-2” Fan/airflow 0 & 1 test fail 350, Bit 22 & 21

“DU# 2,3, or 4 VALID” “PASS” or “FAIL” or“N/A”

“DU# 2,3, or 4 or “FAIL 1” Fan/airflow 0 test failure See Note 2COOLING FANS or “FAIL 2” Fan/airflow 1 test failure 350, Bit 22

or “FAIL 1-2” Fan/airflow 0 & 1 test fail 350, Bit 21 350,Bit 22 & 21

NOTE 1: PFD wrapback data is set via an ARINC 429 bus.

NOTE 2: PFD, MFD, and EICAS VALIDs are true if the DU data received over the HDLC bus is fresh.



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3.3.4 EVENT CODES (RA SETTING 670)

This page displays non-volatile flight data, referred to as Event Codes (EC). A total of 64flights (each flight is defined as one weight on wheels cycle) can be stored. The 65th flightwill automatically overwrite flight number one’s data. Flights will only be stored when an EChas been recorded.

This maintenance page powers up displaying the EC addresses for the current flight, i.e., ECsfor the upcoming flight will be stored in the displayed addresses. If no recordable eventcodes have occurred during the most recent flight, the last flight with recorded event codeswill be displayed. The maximum number of ECs stored per flight is 48. Note that only 8 ECscan be displayed on the screen at once. To view the previous 8 ECs, (“FULL/WX”) and toview the next 8 ECs, use DC-550 “ET,” or press the NAV PB on.

X X X X : X X

Y Y Y Y : Y Y

Y Y Y Y : Y Y

Y Y Y Y : Y Y

Y Y Y Y : Y Y

Y Y Y Y : Y Y

Y Y Y Y : Y Y

Y Y Y Y : Y Y

Y Y Y Y : Y Y

E V E N T C O D E S

F L I G H T 1

E C X X C C C C W O W

E C X X C C C C W O W

E C X X C C C C A I R






670 RA

AD-51502@

D C 5 5 0 P U S H B U T T O N S

G P S D / T T GE TN A V

I N C R E C O D EN E X T F L TP R E V F L T

Figure 3.3.10. Event Code Page



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The flights are numbered from 1 - 64 (Flight 1 is shown in the example) and the most recentflight number will be displayed in green. If a ‘weight on wheels’ cycle has occurred, the startof flight timestamp will be displayed (as indicated by ‘XXXXXXX’), otherwise 00000:00 will bedisplayed for the timestamp. See Table 4.1 and 4.2 for Event Code Descriptons.

The Event Codes are recorded for Autopilot Monitors, Flight Director, EFIS, and EICAS(Table 4.1) and Autopilot Inner Loop Monitors (Table 4.2).

The information displayed on the event code page includes:

PARAMETER SYMBOL COLOR

Event Code Prefix EC White

Event Code Number xx Magenta

Event Code cccc Cyan

Where EC Occurred WOW/AIR Cyan

Time of EC HHHHH:MM Cyan

Start of Flight XXXXX:XX Cyan

Display Controller (DC-550) Push Buttons:

DC-550 PUSHBUTTON FUNCTION

“NAV” View previous 8 event codes (current flight)

“ET” View next 8 event codes (current flight)

GSPD/TTG Increments 1 flight (48 event codes)



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3.3.5 POWER AND INTERNAL TEMPERATURE PAGE (RA SETTING 680)

This page displays the power measured by the IC-600, main power and servo power. Theaverage power while the aircraft was in flight for the main and servo power is displayed toprovide power degradation monitoring (approximately 3 hours of flight).

2 8 . 0 V

2 7 . 5

2 8 . 0 V

2 6 . 0

0 3 8 . 0

0 0 1 . 0

1 1 0 . 0

P O W E R / I N T E R N A L T E M P

M A I N P W R 2 8 V D C

M A I N P W R A V G

S E R V O P W R 2 8 V D C

S E R V O P W R A V G

I N T E R N A L T E M P

L O W E S T E N C O U N T

H I G H E S T E N C O U N T

680 RAAD-51503@

Figure 3.3.11. Power/Internal Temperature Page

CAUTION: DO NOT INSTALL ANOTHER UNIT UNTIL THE DISCREPANCIES WITHTHE POWER BUS ARE RESOLVED.

The power bus in question should be verifed per the Installation EB, and any discrepanciesshould be resolved.

The Installation Bulletin for the EMB-145 provides information for the IC-600 powerrequirements . If a value is found to be outside the limits (32.0<Value<27.5 volts), denoted by“1”, the following actions should be taken:

Lines 3 and 5, column 26:

CONDITION STATUS DISPLAYED COLOR

Value within limits “V” Green

Value outside limits “I” Red

Not installed “N/A” White

The INTERNAL TEMP of the IC-600 is displayed in Celsius. The lowest and highesttemperatures since the unit was initialized by factory personnel is displayed.



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3.3.6 AHRS DATA (RA SETTING 690)

This page shows the pitch, roll, and heading data from the primary and secondary AHRSARINC 429 data bus inputs. The data is displayed in degrees from 0 to +/- 180.0. The busstatus is displayed to the right of the digital readout as follows:

LABEL STATUS DISPLAYED STATUS COLOR

Valid V Green

Test V Green

NCD I Red

Fail I Red

Whenever the staus is displayed as “I”, the following actions should be taken:

• Enter GMT on the cross-side IC-600 and observe the parameter in question. For example,if the parameter is question is the Primary Pitch, then the cross-side IC-600 Secondarypitch should be observed.

• If the two separate GMTs agree, then replace the AHRS with the cross-side AHRS andrepeat the GMT. If the status remains displayed as “I”, then the AHRS should be replacedand the discrepant condition documented. Note that the primary pitch for IC-600 No.1 isAHRS No. 1 and the Secondary pitch for IC-600 No. 2 is AHRS No. 2, while the PrimaryPitch for IC-600 No. 2 is AHRS No. 2 and the Secondary Pitch for IC-600 No. 2 is AHRSNo. 1.

• If the two separate GMTs do not agree, verify the ARINC 429 interface to the IC-600s andthe AHRS.

X X X . X V

X X X . X V

X X X . X V

X X X . X V

X X X . X V

X X X . X V

A H R S D A T A

P R I P I T C H

S E C P I T C H

P R I R O L L

S E C R O L L

P R I H D G

S E C H D G

690 RAAD-51504@

Figure 3.3.12. AHRS Data Page



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3.3.7 DC VARIABLES (RA SETTING 700)

X X . X

X X . X

X X . X

X X X . X

D C V A R I A B L E S

R A D A L T B E L

R A D A L T A B V

T U R N K N O B

P I T C H W H E E L

700 RAAD-51505@

V

V

V

V

V

V

V

V

Figure 3.3.13. DC Variables Page

Line 3 contains the Radio Altitude voltage (J2A-50,51)

Line 4 contains the PC-400 Turn knob voltage (J2A-73)

Line 5 contains the PC-400 Pitch Wheel voltage (J2B-68,69)

The Turn Knob and Pitch Wheel on the PC-400 should be rotated in both directions to verifytheir interface with the IC-600.

When an invalid condition is displayed, the wiring interface between the IC-600 and thePC-400 should be verified. Note that the No. 2 IC-600 will normally display N/A for the TurnKnob and Pitch Wheel.

CONDITION STATUS DISPLAYED COLOR

VALID “V” Green

INVALID “I” Red

Not installed “N/A” White



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3.3.8 MICRO AIR DATA COMPUTER (RA SETTING 710)

A I R D A T A

P R I M A D C V A L I D

S E C M A D C V A L I D

P R I B A R O A L T

S E C B A R O A L T

P R I A L T R A T E

S E C A L T R A T E

P R I C A L A I R S P E E D

S E C C A L A I R S P E E D

710 RA

P A S S

F A I L

X X X X X X

X X X X X X

X X X X

X X X X

X X X X

X X X X

AD-51506@

Figure 3.3.14. Air Data Page

The Barometric Altitude can be changed by rotating the Baro Set knob on the PFD BezelControllers. This will verify the interface from the PFD Bezel Controller to the MADC, as wellas the ARINC 429 interface from the MADC to the IC-600.



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Whenever a “FAIL” condition is displayed for all the parameters, the following actions shouldbe taken:

• Verify the appropriate MADC is installed and power is applied.

• Perform GMT on the cross-side IC-600. If all the parameters also show a “FAIL” condition,the suspect MADC should be replaced.

• If the parameters show a “VALID” condition for the cross-side GMT test, then the ARINC429 interface for the suspect MADC should be verified.

Whenever a “FAIL” condition is displayed for only one or some of the parameters, thefollowing actions should be taken:

• The MADC should be replaced with the cross-side MADC. If the parameters now display a“VALID” condition, the suspect MADC should be replaced.

• If the “FAIL” condition remains, verify the pitot/static interface with the suspect MADC.

LABEL PARAMETER ARINC 429 STATUS

“MADC VALID” Discrete #1 270, BIT 13 Pass, Fail, or N/A

“BARO ALTITUDE (ft)” Altitude 204

“ALTITUDE RATE (ft/min)” Altitude Rate 212

“CAL AIRSPEED (KTS)” Calibrated Airspeed 206

“BARO SETTING (IN-HG)” Barometric Setting 235

“TOTAL AIR TEMP (DEG C)” TAT 211



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3.3.9 DC-550 PORT (RA SETTING 720)

X X

X X

X X

X X

X X

X X

Y Y Y Y Y Y Y Y

Z Z

D C - 5 5 0 P O R T

C R S K N O B

H D G K N O B

S P D K N O B

A S E L K N O B

M F D K N O B

E I C A S K N O B

P B 1 - 2 5

S W 1 - 8

720 RAAD-51507@

Figure 3.3.15. DC-550 Port Page



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The various knob data is shown is HEX format. The knobs should be rotated and acorresponding change noted in the HEX data to verify the interface.

DC-550, GC-550, PFD/MFD bezel controllers, and other controller pushbutton activity aresent to the IC-600 via discretes. The following pushbutton (PB 1-25) discretes are received:

PUSHBUTTON LOCATION

FULL/WX DC-550

GSPD/TTG DC-550

ET DC-550

NAV DC-550

FMS DC-550

MFD Bezel Pushbuttons 1-6 MFD Bezel

Master Caution Inst. Panel

Master Warn Inst. Panel

In/Pa PFD Bezel

SPD IAS Mach GC-550

CRS 1 Sync (Note) GC-550

HDG Sync GC-550

NOTE: Copilots side will display CRS 2.

Engaging the pushbutton will change the status from ‘0’ to ‘1’. Disengaging the pushbuttonwill return the status back to ‘0’.

Additional GC-550 functions are sent through the DC-550 controller before being received bythe IC-600.

• * - AP and YD are only brought in on the pilot side IC-600, and hence will only be displayedif the DU is being driven by the pilot’s IC-600.



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The Reversionary Controller switch inputs are sent to the IC-600 via the DC-550. Thefollowing switch(SW 1-8) discretes are received:

SWITCH LOCATION

AHRS Reversion Reversionary Controller

MADC Reversion Reversionary Controller

DAU Channel A Reversion DAU Rev. Controller

DAU Channel B Reversion DAU Rev. Controller

DU #2/#4* PFD Select Reversionary Controller

DU #2/#4* EICAS Select Reversionary Controller

NOTE: * Copilot’s side.

Engaging the switch will change the status from ‘0’ to ‘1’. Disengaging the pushbutton willreturn the status back to ‘0’.

Whenever a pushbutton or switch is activated and the status does not change from ‘0’ to ‘1’,the following actions should be followed.

• Verify the pushbutton or switch activity on the cross-side IC-600 GMT.

• If the cross-side GMT functions correctly, verify the appropriate wiring interface to theDC-550 or IC-600. If the wiring interface is verified to be correct, replace the DC-550 withthe cross-side DC-550 and verify the switch/pushbutton operation. If it now functionsreplace the suspect DC-550

• If the cross-side GMT also does not correctly function, verify proper operation of the switchor pushbutton in question.



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3.3.10 WEATHER RADAR (RA SETTING 730)

P A S S

W E A T H E R R A D A R

W X S T A T U S

AD-51508@


F U L L / W X S E L / N O T S E L

730 RA

Figure 3.3.16. Weather Radar Page

The Weather Radar status field will display the following:

FAIL is displayed if either the WX Status, WX R/T, WX Mode/Range, or WX Tilt, datamessages on the WX control bus are invalid.

PASS is displayed if the above data messages are all received valid.

NOT SEL is displayed if the WX is not selected for display on the PFD.

Whenever a FAIL condition is observed, verify the wiring interface between the RadarController and the IC-600.



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3.3.11 DISCRETES 1 (RA SETTING 740)

2 8 V

2 8 V

2 8 V

2 8 V

2 8 V

O P N

O P N

O P N

O P N

O P N

D I S C R E T E S 1

J 1 A - 0 7 6 M N T R M D N M

J 1 A - 0 7 6 M N T R M D N A P

J 1 A - 0 7 7 M N T R M U P M

J 1 A - 0 7 7 M N T R M U P A P

J 1 A - 0 7 8 M N T R M E N A B L

J 1 A - 0 7 9 R E S E R V E D

J 1 B - 0 1 4 N 2 O V R R I D E

J 1 B - 2 1 L A M P T E S T I N

J 1 B - 0 2 2 H D G S E L P B

J 1 B - 0 2 3 R E S E R V E D

740 RAAD-51509@

Figure 3.3.17. Discretes 1

IC-600 PIN# FUNCTON SOURCE

J1A-076 Manual Trim Down Manual Pitch Trim Switch

J1A-077 Manual Trim Up Manual Pitch Trim Switch

J1A-078 Reserved

J1A-079 Reserved

J1B-014 N2 Override Maintenance Panel

J1B-021 Lamp Test Input

J1B-022 Heading F/D Pushbutton

J1B-023 Reserved GC-550, 11J1-11



Page 7-43Nov 1/1996


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O P N

O P N

O P N

O P N

O P N

O P N

O P N

O P N

2 8 V

O P N

D I S C R E T E S 2

J 1 B - 0 2 4 N A V S E L P B

J 1 B - 0 2 5 A P P S E L P B

J 1 B - 0 2 6 A L T S E L P B

J 1 B - 0 2 7 S P D S E L P B

J 1 B - 0 2 8 F L C S E L P B

J 1 B - 0 2 9 G A S E L E C T

J 1 B - 0 3 0 V S S E L P B

J 1 B - 0 4 0 R E S E R V E D

J 1 B - 0 4 1 R E S E R V E D

J 1 B - 0 4 4 M S T R W A R N

750 RAAD-51510@


IC-600 PIN# FUNCTION SOURCE

J1B-024 NAV F/D Pushbutton GC-550, 11J1-13

J1B-025 APR F/D Pushbutton GC-550, 11J1-14

J1B-026 ALT F/D Pushbutton GC-550, 11J1-15

J1B-027 SPD F/D Pushbutton GC-550, 11J1-16

J1B-028 FLC F/D Pushbutton GC-550, 11J1-17

J1B-029 GA F/D Pushbutton Throttle Levers

J1B-030 VS F/D Pushbutton GC-550, 11J1-19

J1B-040 Reserved

J1B-041 Reserved

J1B-044 AWC Master Warning Aural Warning Computer



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2 8 V

G N D

2 8 V

5 V

O P N

O P N

2 8 V

O P N

2 8 V

O P N

D I S C R E T E S 3

J 1 B - 0 4 5 J S T K R I G H T

J 1 B - 0 4 8 F D S E L O N

J 1 B - 0 4 9 R E S E R V E D

J 1 B - 0 5 2 R E S E R V E D

J 1 B - 0 7 6 J S T K U P

J 1 B - 0 7 7 J S T K D O W N

J 1 B - 0 7 8 S T I C K S H A K E R

J 1 B - 0 7 9 J S T K L E F T

J 2 A - 0 4 9 R A D A L T V A L

J 2 A - 0 5 4 L S I D E S E L

760 RAAD-51511@


IC-600 PIN # FUNCTION SOURCE

J1B-045 Joystick Input (Right) Optional Joystick

J1B-048 F/D 1 ON Pushbutton GC-550, 11J1-7

J1B-049 Reserved

J1B-052 Reserved

J1B-076 Joystick Input (Up) Optional Joystick

J1B-077 Joystick Input (Down) Optional Joystick

J1B-078 Stall Warning Stall Protection Computer

J1B-079 Joystick Input (Left) Optional Joystick

J2A-049 Radio Alt 28V Valid RAD ALT, 20J1-Y

J2A-054 Left Side Select Aircraft Wiring



Page 7-45Nov 1/1996


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G N D

O P N

O P N

O P N

G N D

G N D

O P N

O P N

O P N

G N D

D I S C R E T E S 4

J 2 A - 0 5 9 R E S E R V E D

J 2 A - 0 6 0 R E S E R V E D

J 2 A - 0 6 1 G E A R C M D

J 2 A - 0 6 2 S P D B R A K E

J 2 A - 0 6 3 S G R E V X S

J 2 A - 0 6 4 W O W

J 2 A - 0 6 5 R E S E R V E D

J 2 A - 0 6 6 E N G / M E T R I C

J 2 A - 0 6 7 R E S E R V E D

J 2 A - 0 6 8 F M S C O N F I G

770 RAAD-51512@



J2A-059 Reserved

J2A-060 Reserved

J2A-061 Ldg Gear Command Landing Gear Lever

J2A-062 Speed Brakes Deployed Speed Brakes

J2A-063 Cross-Side SG Rev SG Reversion Switch

J2A-064 Weight on Wheels (WOW) Squat Switch

J2A-065 Memory Program Enable Switch Memory Program

J2A-066 System Config Pins, Ref Aircraft Wiring



NOTE: Changing WOW, weight on wheels, to “OPN” will result in the maintenance testbeing exited.



Page 7-46Nov 1/1996


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G N D

O P N

O P N

O P N

O P N

G N D

O P N

G N D

O P N

O P N

D I S C R E T E S 5

J 2 A - 0 6 9 A D F C O N F I G

J 2 A - 0 7 0 F D B A R S

J 2 A - 0 7 1 C A T I I E N

J 2 A - 0 7 2 A / C C O N F 1

J 2 A - 0 7 3 A / C C O N F 2

J 2 A - 0 7 4 A / C C O N F 3

J 2 A - 0 7 5 A / C C O N F 4

J 2 A - 0 7 6 A / C C O N F 5

J 2 A - 0 7 7 T E S T I D 1

J 2 A - 0 7 8 T E S T I D 2

780 RAAD-51513@



J2A-069 Sys Config Pins, Ref Aircraft Wiring



J2A-072 Aircraft Configuration Pins Aircraft Wiring





J2A-077 Test Idents Used by Repair andTesting Facilities

Should always read OPN wheninstalled in aircraft


Should always read OPN wheninstalled in aircraft.



Page 7-47Nov 1/1996


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O P N

O P N

O P N

2 8 V

G N D

2 8 V

G N D

2 8 V

G N D

G N D

D I S C R E T E S 6

J 2 A - 0 7 9 T E S T I D 3

J 2 A - 0 8 0 T E S T I D 4

J 2 B - 0 2 2 R E S E R V E D

J 2 B - 0 4 5 R E S E R V E D

J 2 B - 0 4 8 S G R E V O S

J 2 B - 0 4 9 R E S E R V E D

J 2 B - 0 5 5 C P L S E L P B

J 2 B - 0 5 6 L O W B A N K P B

J 2 B - 0 5 7 R E S E R V E D

J 2 B - 0 5 8 R E S E R V E D

790 RAAD-51514@







J2B-022 Reserved

J2B-045 Secondary Trim Up Secondary Trim Switch

J2B-048 SG Reversion SG Reversion Switch

J2B-049 Reserved

J2B-055 Couple Push Button Input GC-550, 11J1-10

J2B-056 Low Bank PB Input GC-550, 11J1-12

J2B-057 Reserved

J2B-058 Reserved



Page 7-48Nov 1/1996


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D I S C R E T E S 7

J 2 B - 0 6 6 M N T R I M U P

J 2 B - 0 6 7 M N T R I M D N

J 2 B - 0 7 0 T C S M A I N

J 2 B - 0 7 0 T C S A P

J 2 B - 0 7 1 R E S E R V E D

J 2 B - 0 7 2 R E S E R V E D

J 2 B - 0 7 5 T R N K N O B

J 2 B - 0 7 6 A P D I S C M

J 2 B - 0 7 6 A P D I S C A P

J 2 B - 0 8 3 Y D E N G P B

O P N

O P N

O P N

O P N

O P N

2 8 V

O P N

2 8 V

2 8 V

G N D

800 RAAD-51515@



J2B-066 Manual Trim Up Trim Switch

J2B-067 Manual Trim Dn Trim Switch

J2B-070 TCS 28V Input MAIN CPU TCS Switch

J2B-070 TCS 28V Input AP CPU TCS Switch

J2B-071 Reserved

J2B-072 Reserved

J2B-075 Turn Knob Out of Detent

J2B-076 M Disc Input to main CPU M disc switches; 28V when notpushed, OPEN when pushed

J2B-076 AP Disc Input to main CPU AP disc switches; 28V when notpushed, OPEN when pushed

J2B-083 Yaw Engage Select PB In GC-550, 11J1-9



Page 7-49Nov 1/1996


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D I S C R E T E S 8

J 2 B - 0 8 5 A P E N G P B

J 2 B - 0 9 0 A P D I S C M

J 2 B - 0 9 0 A P D I S C A P

J 2 B - 0 9 1 M S T R C A U T

J 2 B - 0 9 4 R S I D E S E L

G N D

2 8 V

2 8 V

O P N

O P N

810 RAAD-51516@



J2B-085 AP Engage Select PB Input GC-550, 11J1-8

J2B-090 AP Disc 28V In Main CPU IC-600, J2B-076

J2B-090 AP Disc 28V In AP CPU IC-600, J2B-076

J2B-091 AWC Master Caution Aural Warning Computer

J2B-094 Right Side Select Aircraft Wiring



Page 7-50Nov 1/1996


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D I S C R E T E S 9

A P C L U T C H M A I N

A P C L U T C H A P

Y D C L U T C H M A I N

Y D C L U T C H A P

A P E N G E N * M A I N

A P E N G E N * A P

Y D E N G E N * M A I N

Y D E N G E N * A P

A P D I S C O U T E A

E L E V S E R V O D I R

O P N

O P N

O P N

O P N

L O

L O

L O

L O

L O

L O

820 RAAD-51517@


This page displays IC-600 internal discretes.

The Pitch, Roll, and Elevator Servo Clutches are all activated from a single discrete output.This output is wrapped around to both processors’ input discretes internally. The autopilotprocessor then passes its discrete back to the main processor for display on the maintenancepage.



Page 7-51Nov 1/1996


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3.3.20 LAMP TEST 9 (RA SETTING 830)

L A M P S T E S T

A N N U N C I A T E D L A M P M S T R W R N


G P S D / T T G I N C R T E S TE T D E C R T E S T AD-51518@

830 RA

Figure 3.3.26. Lamp Test



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The lamp driver test is executed by sequencing through the lamp test sequence shown below.

LAMP BEING TESTED TEST SEQUENCE

PILOT’S SIDE:Cockpit Lamps MSTR WRN, MSTR CAU

GC-550 FD1, HDG, NAV, APR, BNK, AP, CPL LEFT,YD, SPD, FLC, VS, ALT.

COPILOT’S SIDE:Cockpit Lamps MSTR WRN, MSTR CAU.

GC-550 FD2, HDG, NAV, APR, BNK, AP, CPLRIGHT, YD, SPD, FLC, VS, ALT.

The GSPD/TTG pushbutton increments to the desired lamp driver test, and the ETpushbutton is decrementa to the desired lamp driver test. The “ANNUNCIATED LAMP” linecorresponds to the lamp being tested.



Page 7-53Nov 1/1996


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3.3.21 ARINC 1 (RA SETTING 840)

A R I N C 1

0 J 2 A - 0 2 1 L R N P

J 2 A - 0 2 2

1 J 2 B - 0 1 1 A P

J 2 B - 0 1 2

2 J 2 A - 0 2 3 C M C

J 2 A - 0 2 4

3 J 2 B - 0 2 3 A D C P

J 2 B - 0 2 4

V A L D A T A

A P C P U

U N U S E D

N C D D A T A

840 RAAD-51519@

Figure 3.3.27. ARINC #1

Because equipment installed on the aircraft may vary as identified by the configurationdiscretes, the ARINC 429 channels shown must be tailored accordingly. For instance, therewould be no channel descriptions next to the pin and the status would be “UNUSED” if theequipment were not installed.

STATUS DISPLAYED COLOR

“VAL DATA” Green

“TST DATA” Amber

“NCD DATA” Amber

“INV DATA” Red

“NO DATA” Red

“UNUSED” Dim white, displayed if equipment is not installed

“NOT SEL” Amber, displayed for channels 0, 2, 3 if theselected source is secondary

N/A N/A for A/P ARINC is normal for IC-600 #2



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A R I N C 2

4 J 2 A - 0 2 9 A H R S P

J 2 A - 0 3 0

5 J 2 B - 0 2 9 T C A S

J 2 B - 0 3 0

6 J 2 A - 0 3 1 R D A U S

J 2 A - 0 3 2

7 J 2 B - 0 3 1 L D A U S

J 2 B - 0 3 2

U N U S E D

V A L D A T A

N O D A T A

N C D D A T A

AD-51520@

850 RA



STATUS DISPLAYEED COLOR



“INV DATA” Red


“NO DATA” Red


“NOT SEL” Amber, displayed for channels 6 and 7 if the selected sourceis secondary



Page 7-55Nov 1/1996


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V A L D A T A

V A L D A T A

V A L D A T A

U N U S E D

860 RAAD-51521@

A R I N C 3

8 J 2 A - 0 4 1 R D A U P

J 2 A - 0 4 2

9 J 2 B - 0 8 1 L D A U P

J 2 B - 0 8 2

1 0 J 2 A - 0 4 3 A D C S

J 2 A - 0 4 4

1 1 J 2 A - 0 8 3 A H R S S

J 2 A - 0 8 4






“INV DATA” Red


“NO DATA” Red


“NOT SEL” Amber, displayed for channels 9 and 10 if theselected source is secondary



Page 7-56Nov 1/1996


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U N U S E D

V A L D A T A

U N U S E D

U N U S E D

870 RAAD-51522@

A R I N C 4

1 2 J 2 A - 0 4 0 S P A R E

J 2 A - 0 2 6

1 3 J 2 B - 0 1 3 W N D S H R

J 2 B - 0 2 6

1 4 J 2 A - 0 2 7 S P A R E

J 2 A - 0 2 8

1 5 J 2 B - 0 2 7 O S P F D

J 2 B - 0 2 8






“INV DATA” Red


“NO DATA” Red


“NOT SEL” Amber, displayed for channels 12 and 14 if theselected source is secondary



Page 7-57Nov 1/1996


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3.3.25 RADIO SYSTEM BUS (RA SETTING 880)

R A D I O S Y S B U S

S R N P R I

S R N S E C

D M E P R I

D M E S E C

V A L D A T A

V A L D A T A

V A L D A T A

V A L D A T A

880 RAAD-51523@

Figure 3.3.31. Radio System Bus

Equipment installed on the aircraft may vary as identified by the configuration discretes for theSRN and data provided via the RSB for the DME. Therefore, if the equipment is not installedthe RADIO SYS BUS page will display “UNUSED” for that equipment in the status column.

STATUS DISPLAYED COLOR DESCRIPTION


“INV DATA” Red

“NO DATA” Red

“UNUSED” Dim white 1. For SRN check config discretes

2. For DME check configuration via RSB



Page 7-58Nov 1/1996


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3.3.26 Checklist Loading (RA Setting 890)

The IC-600 supports the loading of user defined checklists via RS-232 interface with theUser-Friendly Checklist System (UFCS), HPN (TBD).

The UFCS will interface with the IC-600 through the BOOT software.

The IC-600 will program the checklist via commands from the UFCS. The UFCS will createthe data exchange and user interface.

890 RA

AD-51524@

C H E C K L I S T L O A D I N G

P R O G R A M M A B L E C H E C K L I S T

E Q U I P M E N T I S R E Q U I R E D

S C R E E N W I L L B L A N K

C Y C L E I C - 6 0 0 C I R C U I T

B R E A K E R T O R E C O V E R


N A V E N T E R C K L S T

Figure 3.3.32. Check Loading Page



Page 7-59Nov 1/1996


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3.3.27 SYSTEM CONFIGURATION (RA SETTING 900)

S Y S C O N F I G

A / C C O N F I G

E N G L I S H / M E T R I C

F M S

A D F

F D C O M M A N D B A R S

C A T I I A P R

E M B - 1 4 5

E N G L I S H

N O T I N S T

S I N G L E

S I N G C U E

D I S A B L E D

900 RAAD-51525@

Figure 3.3.33. System Configuration

3.3.28 BLANK PAGE (RA SETTING 910)

B L A N K P A G E

B L A N K P A G E

B L A N K P A G E

B L A N K P A G E

B L A N K P A G E

B L A N K P A G E

B L A N K P A G E

B L A N K P A G E

B L A N K P A G E

B L A N K P A G E

B L A N K P A G E

B L A N K P A G E

910 RAAD-51526@

Figure 3.3.34. Blank Page



Page 7-60Nov 1/1996


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4. EVENT CODES

NOTE 1: If an Event Code No. is displayed and it is not listed in this document, Event Codes,please contact your local Customer Support Engineer for further instruction.

NOTE 2: Please write the Event Code No. on your exchange or repair transaction. Pleaseprovide details regarding maintenance actions required to fix system anomalies.

There are two categories of IC-600 failure and associated codes:

(a) Power-up failures - These result in a latched failure which can only be reset by cycling the 28Vdc circuit breaker. Power-up failures are identified as Event Codes 0001-0199 and all 90XXcodes.

(b) Continuous monitored failures - These result in AP/YD disengage with an associated EventCode, all 02XX and 91XX codes. If one of these monitors trips, such as in overpowering theautopilot, and the AP FAIL annunciation is displayed on the EFIS/EICAS, the failure can bereset by holding the AP DISC button for 2 second. Resetting the IC-600 circuit breaker willalso reset the computer, however, the preferred method is the AP disconnect button. Not all ofthe 02XX or 91XX codes will result in the AP failing. They may result in the AP disconnecting,for example, a code 0207 - Primary Pitch invalid.

Event codes 03XX and 04XX are codes which are logged but do not prevent AP engagement.

CAUTION: IT IS IMPORTANT TO NOTE THAT AT NO TIME SHOULD MAINTENANCEPERSONNEL ATTEMPT TO REPLACE A HONEYWELL LRU BASED ONEVENT CODES ALONE. IN MOST CASES, THE RECORDED EVENT CODESSHOULD BE USED AS A TOOL ALONG WITH TRADITIONAL GROUNDTESTING AND OTHER STANDARD MAINTENANCE PRACTICES.

Table 4.1. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes(Power Up Event Codes That Prevent Autopilot Engagement)

MODULE EVENT CODE # DESCRIPTION

Engtest 0001 TCS engaged during power up test

Engtest 0002 GA engaged during power up test

Engtest 0003 Heartbeat latch is not reset before executing Engtest

Engtest 0004 Software cannot disengage the autopilot

Engtest 0005 Software cannot disengage the yaw damper

Engtest 0006 AP CPU Engtst acknowledge timeout error

Engtest 0007 Heartbeat latch cannot disengage the autopilot

Engtest 0008 Heartbeat latch cannot enable autopilot and yaw damper

Engtest 0009 AP valid discrete cannot disable the autopilot

Engtest 0010 AP valid discrete cannot enable the AP engage enable discrete

Engtest 0011 YD valid discrete cannot disable the yaw damper

Engtest 0012 YD valid discrete cannot enable the YD engage enable discrete

Engtest 0013 AP CPU Engtst acknowledge timeout error

Srvtst 0014 Power supply flipflop is not reset after clock pulse



Page 7-61Nov 1/1996


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Srvtst 0015 AP elevator servo tach moved when the AP servo valid wasinvalid and the YD servo valid was valid

Srvtst 0016 AP aileron servo tach moved when the AP servo valid was invalidand the YD servo valid was valid

Srvtst 0017 AP rudder servo tach moved when the AP servo valid was validand the YD servo valid was invalid

Srvtst 0018 Trim end-arounds enabled when the AP servo valid was invalidand the YD servo valid was valid

Srvtst 0019 Elevator servo tach moved when the AP servo valid was invalidand the YD servo valid was valid

Srvtst 0020 Aileron servo tach moved when the AP servo valid was invalidand the YD servo valid was valid

Srvtst 0021 Rudder servo tach moved when the AP servo valid was valid andthe YD servo valid was invalid

Srvtst 0022 Excessive AP elevator servo current when the AP servo validwas invalid and the YD servo valid was valid

Srvtst 0023 Excessive AP aileron servo current when the AP servo valid wasinvalid and the YD servo valid was valid

Srvtst 0024 Excessive AP rudder servo current when the AP servo valid wasvalid and the YD servo valid was invalid

Srvtst 0025 Excessive elevator servo current when the AP servo valid wasinvalid and the YD servo valid was valid

Srvtst 0026 Excessive aileron servo current when the AP servo valid wasinvalid and the YD servo valid was valid

Srvtst 0027 Excessive rudder servo current when the AP servo valid wasvalid and the YD servo valid was invalid

Srvtst 0028 Elevator servo direction discrete indicates wrong currentdirection

Srvtst 0029 Elevator trim end-arounds are not indicating that elevator istrimming up

Srvtst 0030 AP elevator servo tach moved too far after 10 milliseconds

Srvtst 0031 AP aileron servo tach moved too far after 10 milliseconds

Srvtst 0032 AP rudder servo tach moved too far after 10 milliseconds

Srvtst 0033 Elevator servo tach moved too far after 10 milliseconds

Srvtst 0034 Aileron servo tach moved too far after 10 milliseconds

Srvtst 0035 Rudder servo tach moved too far after 10 milliseconds

Srvtst 0036 AP elevator current not high enough after 10 milliseconds

NOTE: Occurs when servos are not connected

Srvtst 0037 AP aileron current not high enough after 10 milliseconds



Page 7-62Nov 1/1996


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Srvtst 0038 AP rudder current not high enough after 10 milliseconds

Srvtst 0039 Elevator current not high enough after 10 milliseconds

Srvtst 0040 aileron current not high enough after 10 milliseconds

Srvtst 0041 Rudder current not high enough after 10 milliseconds

Srvtst 0042 AP elevator servo tach did not move enough after 60milliseconds

Srvtst 0043 AP aileron servo tach did not move enough after 60 milliseconds

Srvtst 0044 AP rudder servo tach did not move enough after 60 milliseconds

Srvtst 0045 Elevator servo tach did not move enough after 60 milliseconds

Srvtst 0046 Aileron servo tach did not move enough after 60 milliseconds

Srvtst 0047 Rudder servo tach did not move enough after 60 milliseconds

Srvtst 0048 Excessive AP elevator current after 60 milliseconds

Srvtst 0049 Excessive AP aileron current after 60 milliseconds

Srvtst 0050 Excessive AP rudder current after 60 milliseconds

Srvtst 0051 Excessive elevator current after 60 milliseconds

Srvtst 0052 Excessive aileron current after 60 milliseconds

Srvtst 0053 Excessive rudder current after 60 milliseconds

Srvtst 0054 Elevator servo direction discrete indicates wrong currentdirection

Srvtst 0055 Elevator trim end-arounds are not indicating elevator is trimmingdown

Srvtst 0056 AP elevator current too high after 10 milliseconds of polarityreversal

Srvtst 0057 AP aileron current too high after 10 milliseconds of polarityreversal

Srvtst 0058 AP rudder current too high after 10 milliseconds of polarityreversal

Srvtst 0059 Elevator current too high after 10 milliseconds of polarity reversal

Srvtst 0060 Aileron current too high after 10 milliseconds of polarity reversal

Srvtst 0061 Rudder current too high after 10 milliseconds of polarity reversal

Srvtst 0062 AP elevator servo tach did not return to zero 25 millisecondsafter polarity reversal

Srvtst 0063 AP aileron servo tach did not return to zero 25 milliseconds afterpolarity reversal

Srvtst 0064 AP rudder servo tach did not return to zero 25 milliseconds afterpolarity reversal



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Srvtst 0065 Elevator servo tach did not return to zero 35 milliseconds afterpolarity reversal

Srvtst 0066 Aileron servo tach did not return to zero 35 milliseconds afterpolarity reversal

Srvtst 0067 Rudder servo tach did not return to zero 35 milliseconds afterpolarity reversal

Svtsthb 0068 Power supply flipflop is not reset at power up

Svtsthb 0069 Power supply flipflop is not set after clock pulse

Svtsthb 0070 Heartbeat flipflop is not reset at power up

Svtsthb 0071 AP elevator servo tach moved with the heartbeat latchdisengaged

Svtsthb 0072 AP aileron servo tach moved with the heartbeat latch disengaged

Svtsthb 0073 AP rudder servo tach moved with the heartbeat latch disengaged

Svtsthb 0074 Elevator trim end-arounds enabled with the heartbeat latchdisengaged

Svtsthb 0075 Elevator servo tach moved with the heartbeat latch disengaged

Svtsthb 0076 Aileron servo tach moved with the heartbeat latch disengaged

Svtsthb 0077 Rudder servo tach moved with the heartbeat latch disengaged

Srvtst 0078 Heartbeat flipflop is not set before executing Srvtst

Svtstps 0079 Heartbeat flipflop is not set before executing Svtstps

Svtstph 0080 Power supply flipflop is not reset before executing Svtstps

Svtstps 0081 AP elevator servo tach moved with the PS flipflop disengaged

Svtstps 0082 AP aileron servo tach moved with the PS flipflop disengaged

Svtstps 0083 AP rudder servo tach moved with the PS flipflop disengaged

Svtstps 0084 Elevator trim end-arounds enabled with the PS flipflopdisengaged

Svtstps 0085 Elevator servo tach moved with the ps flipflop disengaged

Svtstps 0086 Aileron servo tach moved with the ps flipflop disengaged

Svtstps 0087 Rudder servo tach moved with the ps flipflop disengaged

Svtstps 0088 AP CPU srvtstact acknowledge timeout error

Svtstps 0089 AP CPU srvtstact acknowledge timeout error, used in twodifferent parts of the module, at the beginning and at the end

Srvtst 0090 AP CPU srvtstact acknowledge timeout error

Pstest 0091 Power supply flipflop was not reset before executing Pstest

Pstest 0092 Undervoltage power supply test did not reset the 80960 CPU

Pstest 0093 Power supply undervoltage test did not reset the power supplyflipflop



Page 7-64Nov 1/1996


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Pstest 0094 Power supply flipflop is not reset after clock pulse

Pstest 0095 Overvoltage power supply test did not reset the 80960 CPU

Pstest 0096 Power supply overvoltage test did not reset the power supplyflipflop

Ap_powerup 0097 (WOW) Arinc 429 timeout error, no communication from the80188 CPU

Hbmtest 0098 Heartbeat flipflop is not set before executing Hbmtest

Ap_powerup 0099 Arinc 429 timeout error, a_pwruprdy was not set by the 80188CPU

Ap_powerup 0100 (WOW) Arinc 429 timeout error, no communication from the80188 CPU

Ap_powerup 0101 Power supply flipflop is not reset after clock pulse

Ap_powerup 0102 Heartbeat flipflop is not reset after clock pulse

Srvtst 0103 Elevator current is too low after 60 milliseconds

Srvtst 0104 Aileron current is too low after 60 milliseconds

Srvtst 0105 Rudder current is too low after 60 milliseconds

Srvtst 0106 AP elevator current too high after 10 milliseconds of polarityreversal

Srvtst 0107 AP aileron current too high after 10 milliseconds of polarityreversal

Srvtst 0108 AP rudder current too high after 10 milliseconds of polarityreversal

Hbmtest 0109 Power supply flipflop was not reset before executing Hbmtest

Hbmtest 0110 Power supply flipflop reset after heartbeat monitor failure

Hbmtest 0111 Heartbeat flipflop did not reset after heartbeat failure

Hbmtest 0112 Heartbeat flipflop did not reset after clock pulse

Hbmtest 0113 Power supply flipflop reset after heartbeat monitor failure

Hbmtest 0114 Heartbeat flipflop did not reset after heartbeat monitor failure

Hbmtest 0115 Heartbeat flipflop did not reset after clock pulse

Hbmtest 0116 Two quick writes did not trip the heartbeat monitor

Hbmtest 0117 Heartbeat monitor did not trip after not updating for 90milliseconds

Ap_powerup 0118 Power supply flipflop is not reset after clock pulse

Ap_powerup 0119 Heartbeat flipflop did not reset after clock pulse

Engtest 0120 AP disconnect active during power up test

Engtest 0121 YD select button is pushed or the AP controller is not installed

Engtest 0122 AP select button is pushed during power up test



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Pkdet_tst_pu/c 0123 Primary peak detect initially failed

Pkdet_tst_pu/c 0124 Primary peak detect cannot be failed

Pkdet_tst_pu/c 0125 Primary peak detect failed to go valid

Pkdet_tst_pu/c 0126 Secondary peak detect initially failed

Pkdet_tst_pu/c 0127 Secondary peak detect cannot be failed

Pkdet_tst_pu/c 0128 Secondary peak detect failed to go valid

Errlgc 0129 Aircraft ID mismatch

Pkdet_tst_c 0130 AC reference failed to go valid before the timeout

Pkdet_tst_c 0131 Peak detect test executed an undefined state

Srvtst 0132 AP elevator servo tach moved at the beginning of srvtst

Srvtst 0133 AP aileron servo tach moved at the beginning of srvtst

Srvtst 0134 AP rudder servo tach moved at the beginning of srvtst

Srvtst 0135 Elevator servo tach moved at the beginning of srvtst

Srvtst 0136 Aileron servo tach moved at the beginning of srvtst

Srvtst 0137 Rudder servo tach moved at the beginning of srvtst

Srvtst 0139 Excessive AP pitch current at the beginning of srvtst

Srvtst 0140 Excessive AP roll current at the beginning of srvtst

Srvtst 0141 Excessive AP yaw current at the beginning of srvtst

Srvtst 0142 Excessive pitch current at the beginning of srvtst

Srvtst 0143 Excessive roll current at the beginning of srvtst

Srvtst 0144 Excessive yaw current at the beginning of srvtst

Main 0145 Timer #3 count register failure

Main 0146 Timer #3 interrupt failure

Main 0176 Application Prom CRC test ticket failure



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Table 4.2. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes(Continuous Event Codes That Disengage the AP and YD Autopilot Engagement)


Errlgc 0200 AP monitor ticket failure when the autopilot/yaw damper wasengaged

Errlgc 0201 Ram partition failure when the autopilot/yaw damper wasengaged

Errlgc 0202 Primary mux A/D-D/A wrap around failure when the autopilot/yawdamper was engaged

Errlgc 0203 Secondary mux A/D-D/A wrap around failure when theautopilot/yaw damper was engaged

Errlgc 0204 The AP monitor was not finished in the time allotted when theautopilot/yaw damper was engaged

Errlgc 0205 Continuous test detected a CRC checksum failure when theautopilot/yaw damper was engaged

Errlgc 0206 Rate gyro went invalid when the autopilot/yaw damper wasengaged

Errlgc 0207 Primary pitch VG went invalid when the autopilot/yaw damper wasengaged

Errlgc 0208 Secondary pitch VG went invalid when the autopilot/yaw damperwas engaged

Errlgc 0209 Primary roll VG went invalid when the autopilot/yaw damper wasengaged

Errlgc 0210 Secondary roll VG went invalid when the autopilot/yaw damperwas engaged

Errlgc 0211 The 80188’s AP valid went invalid when the autopilot/yaw damperwas engaged

Errlgc 0212 AP Arinc 429 went invalid when the autopilot/yaw damper wasengaged

Errlgc 0213 Flight director data went invalid when the autopilot/yaw damperwas engaged

Errlgc 0214 Stick shaker went invalid when the autopilot/yaw damper wasengaged

Errlgc 0215 Pgyro’s pitch attitude monitor tripped when the autopilot/yawdamper was engaged

Errlgc 0216 Priatt’s pitch attitude monitor tripped when the autopilot/yawdamper was engaged

Errlgc 0217 Rgyro’s roll attitude monitor tripped when the autopilot/yawdamper was engaged

Errlgc 0218 Priatt’s roll attitude monitor tripped when the autopilot/yawdamper was engaged



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Table 4.2. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes(Continuous Event Codes That Disengage the AP and YD Autopilot Engagement)


Errlgc 0219 Logic’s AP disconnect monitor tripped when the autopilot/yawdamper was engaged

Errlgc 0220 Logic’s YD disconnect monitor tripped when the autopilot/yawdamper was engaged

Errlgc 0221 Pmodel’s pitch servo monitor tripped when the autopilot/yawdamper was engaged

Errlgc 0222 Rmodel’s roll servo monitor tripped when the autopilot/yawdamper was engaged

Errlgc 0223 Ymodel’s yaw servo monitor tripped when the autopilot/yawdamper was engaged

Errlgc 0224 Accmon’s vertical acceleration monitor tripped when theautopilot/yaw damper was engaged

Errlgc 0225 Rolmon’s roll rate monitor tripped when the autopilot/yaw damperwas engaged

Errlgc 0226 Tmrun’s trim runaway monitor tripped when the autopilot/yawdamper was engaged

Errlgc 0227 Autopilot monitor input ram corruption

Errlgc 0228 Autopilot monitor output ram corruption

Errlgc 0229 Flight director went invalid when the autopilot/yaw damper wasengaged

Errlgc 0230 YD monitor ticket failure when the autopilot/yaw damper wasengaged

Errlgc 0231 The 80188’s YD valid went invalid when the autopilot/yaw damperwas engaged

Errlgc 0234 Non-AP side flight director went invalid when the autopilot/yawdamper was engaged

Errlgc 0235 Loss of servo power when the autopilot/yaw damper was engaged



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Table 4.3. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes(Power Up Event Codes)


Main 0300 Arinc ram test failure

Ca_init 0301 Local ram test failure, address test

Ca_init 0302 Local ram test failure, marching 1’s

Ca_init 0303 Warm start ram checksum failure

Ca_init 0304 Local ram test failure AA’s, 55”s, FF’s, 00”s

Hdlc_powerup_w 0305 Power up ICB wraparound test failure, timeout failure

Hdlc_powerup_w 0306 Power up ICB wraparound test failure, checksum failed 3consecutive times

Hdlc_powerup_w 0307 Power up ICB wraparound test failure, TX not done or RXnot done

Ap_powerup 0308 Transfer of AP software ID failure, cold start, WOW

Main 0309 RS-422 channel #1 UART TX data timeout error

Main 0310 RS-422 channel #1 UART DMA transfer failure, either theDMA failed or the UART failed to request DAM

Main 0311 RS-422 channel #1 UART TX or RX parity or framing error

Main 0312 RS-422 channel #1 RX data failure, DMA xfer occurred towrong address or ram didn’t accept the data

Main 0313 RS-422 channel #1 UART DMA terminal count failure
















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Table 4.3. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes(Power Up Event Codes)







Main 0336 Power up partition failure in supervisor mode

Main 0337 Power up partition failure in user/essential mode

Main 0338 Power up partition failure in user/non-essential mode

Main 0340 Stored Arinc #1 CRC didn’t agree with the CRC stored ineeprom and the calculated CRC

Main 0341 Stored Arinc #1 CRC didn’t agree with the CRC stored ineeprom

Main 0342 Stored Arinc #1 CRC didn’t agree with the calculated CRC

Main 0344 Stored Arinc #2 CRC didn’t agree with the CRC stored ineeprom and the calculated CRC

Main 0345 Stored Arinc #2 CRC didn’t agree with the CRC stored ineeprom

Main 0346 Stored Arinc #2 CRC didn’t agree with the calculated CRC

Main 0350 Stored Application CRC didn’t agree with the CRC stored ineeprom and the calculated CRC

Main 0351 Stored Application CRC didn’t agree with the CRC stored ineeprom

Main 0352 Stored Application CRC didn’t agree with the calculatedCRC



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Table 4.4. Autopilot Monitors, Flight Director, EFIS, and EICAS Event Codes(Event Codes That Record An Action)


Errlgc 0400 80960 CPU had to reset the 80188 CPU, Arinc 429 timeouterror

Evcode 0401 Event code buffer full

Afcs_pr_sup 0402 Run time reinitialization of ASIC #1 Arinc receivers

Evcode 0403 IC-500 over-temperature has occurred

Evcode 0404 IC-500 Over-temperature Intermittent problem

Xs_input 0405 The non-AP side IC did not acknowledge being the FDmaster in the time allotted

Rev_proc 0406 Invalid DU reversionary mode selected

Tcas_enabled 0407 TCAS flash failure

Pfdwpr 0408 IAS DU wrap around failure

Pfdwpr 0409 ALT DU wrap around failure

Pfdwpr 0410 Pitch DU wrap around failure

Pfdwpr 0411 Roll DU wrap around failure

Ca_init 0412 Warm start bit stuck in the warm start position, or IC resetdue to heartbeat monitor



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Table 4.5. Autopilot and Interloop Monitors Event Codes(Power Up Event Codes That Prevent Autopilot Engagement)


Boot188a 9000 AP power up ram test failure

Engtst 9001 TCS engaged during power up testing

Powrup 9002 Power up ticket failure

Engtst 9003 AP engaged during power up testing

Engtst 9004 YD engaged during power up testing

Engtst 9005 AP or YD engage cannot be enabled

Engtst 9006 AP engage valid cannot disable the autopilot or the APengage valid disables the yaw damper

Engtst 9007 YD engage valid cannot re-enable YD engage enable

Promtest 9008 AP rom CRC checksum failure

Engtst 9009 AP engage valid cannot re-enable AP engage enable

Wraptest 9010 AP A/D power up wrap around failure (mux 0)

Wraptest 9011 AP A/D mux 1 failure

Wraptest 9012 AP-A/D mux 2 failure

B429test 9013 Arinc transmitter #0 not empty

B429test 9014 Timeout error, transmitter interrupt (int3) did not occur

B429test 9015 Timeout error, DMA interrupt did not occur

B429test 9016 Timeout error, receive interrupt error, could be a labeldetection error

B429test 9017 Timeout error, Arinc warp around error

B429test 9018 Arinc 429 DMA wrap around data failure

B429test 9019 Arinc 429 RXO wrap around data failure

B429test 9020 Arinc 429 RX1 wrap around data failure

Btxint 9021 TX interrupt was not cleared by read cycle

Engtst 9022 YD engage valid cannot disable the yaw damper or the YDengage valid disables the autopilot



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Table 4.6. Autopilot and Interloop Monitors Event Codes(Continuous Event Codes That Prevent Autopilot Engagement)


Elogic 9100 VG was invalid when the autopilot/yaw damper wasengaged or when the AP button was pressed

Elogic 9101 AP Arinc 429 was invalid when the autopilot/yaw damperwas engaged or when the AP button was pressed

Elogic 9102 AP ticket failure when the autopilot/yaw damper wasengaged or when the AP button was pressed

Elogic 9103 AP continuous A/D wrap around failure when theautopilot/yaw damper was engaged or when the AP buttonwas pressed

Elogic 9104 AP AC reference failure when the autopilot/yaw damperwas engaged or when the AP button was pressed

Elogic 9105 AP analog ground failure when the autopilot/yaw damperwas engaged or when the AP button was pressed

Elogic 9106 AP detected that the primary processor’s AP ticket wasinvalid when the autopilot/yaw damper was engaged orwhen the AP button was pressed

Elogic 9107 Stack runaway error, exceeded lower ram memoryboundary

Elogic 9108 Event code buffer runaway error, exceeded upper rammemory boundary

Dynrmt 9109 AP continuous ram test failure

Elogic 9110 AP detected that the primary processor’s YD ticket wasinvalid when the autopilot/yaw damper was engaged orwhen the AP button was pressed

Elogic 9111 Background ticket failure

Table 4.7. Autopilot and Interloop Monitors Event Codes Execution Failure


Undefint 9200 Executed an undefined interrupt <type 32

Badint 9201 Executed an undefined interrupt > type 31



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SECTION IISTANDARD GROUND TEST PROCEDURE

1. INTRODUCTION

The purpose of the following tests is to verify functionality of each component and its associatedwiring. Inputs and outputs for each LRU will be tested. IC-600 internal functions such as gainprogramming and monitors will not be exercised as they are verified at power up by softwarechecks.

Prior to testing ensure all sensors and radios are valid.

In many instances the reader will be referenced to Section I, IC-600 Self Test, where componentand interface functionality can be verified.

All values stated in these tests have a ± 10% tolerance.

2. DISCRETE INPUT VERIFICATIONS

The following table lists various discrete inputs to the IC-600. Reference Section I, IC-600 SelfTest, RA Settings, 740 through 810, to observe the input change as the respective switch isactivated or power source (CB, circuit breaker) is pulled.

FUNCTION SWITCH/SOURCE IC-600 RA SETTING

LAMPS TEST LAMP TEST INPUT 830

MASTER WARN AWC MASTER WARNING 830

STICK SHAKER STALL WARNING CB 760

RAD ALT VALID RAD ALT CB 760

WOW SQUAT SWITCH 770

LOWBANK PB BANK SW ON GC-550 790

SG REV REVERSIONARY CONTROLLER 790

TCS TCS SW ON CONT WHEEL 800

GO AROUND GA SW ON THROTTLE 750

TURN KNOB TK ON PC-400 800

AP DISC AP DISC SW ON CONT WHEEL 800 AND 810

YD ENG PB YD BUTTON ON GC-550 800

AP ENG PB AP BUTTON ON GC-550 810

MASTER CAUTION AWC MASTER CAUTION 810



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3. DU-870 (DISPLAY UNIT)

The PFD, MFD, and EICAS DUs are identical except for the Bezel Controller installed on the frontof the units. The displays may be swapped from one position to another to verify a suspecteddefective display or wiring problem.

Step 1: Ensure MFD switch on Reversionary Controller is in “NORM” (normal) position.

Step 2: After applying aircraft power, ensure all displays light up.

Step 3: Check each display for color, clarity, and general overall appearance.

Step 4: Shine a light into each of the ambient light sensors (photo cells) on each display andobserve the display brightness level increases.

Step 5: Refer to Section I, paragraph 3.3.3.1, RA Setting 660. The STATUS column shouldindicate PASS for each item.

4. DC-550 DISPLAY CONTROLLER

The display controller communicates with the IC-600 over a two wire digital bus. This digital buscan be verified by selecting any function on the controller and observing the appropriate responseon the displays.

(a) Refer to Section I, paragraph 3.3.9, RA Setting 720. DC-550 VALID should indicate PASS.

(b) Refer to Section I, paragraph 3.3.9, RA Setting 720. Activate each button/knob/switch on theDC-550. A message corresponding to each activation should be displayed.

(c) Refer to Section I, paragraph 3.3.9, RA Setting 720. Activate the IAS/M, HDG SYNC, andCRS SYNC pushbuttons on the GC-550. A message corresponding to each activation shouldbe displayed.

(d) Refer to Section I, paragraph 3.3.9, RA Setting 720. Activate the Master Warning and MasterCaution acknowledge pushbuttons. A message corresponding to each activation should bedisplayed.

(e) Refer to Section I, paragraph 3.3.9, RA Setting 720. Rotate the CRS, HDG, ALT, and CASknobs on the GC-550 and the M/P RNG knob on the MFD and observe the values change foreach respective knob.



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5. IC-600 INTEGRATED COMPUTER

The IC-600 Integrated Computer provides symbol generator, fault warning, flight director, andautopilot (if installed) functions. Refer to Section I, paragraph 3.2.2, regarding IC-600 self testing.

6. AZ-850 MICRO AIR DATA COMPUTER

(a) Reference Section I, paragraph 3.3.8, RA Setting 710. MADC VALID should display PASSindicating a valid AZ-850 and the integrity of its associated bus wiring to the IC-600.

(b) Reference Section I, paragraph 3.3.21, RA Setting 840. MADC 1 should display VAL DATAindicating a valid AZ-850 and the integrity of its associated bus wiring to the IC-600.

(c) Reference Section I, paragraph 3.3.23, RA Setting 860. MADC 2 should display VAL DATAindicating a valid AZ-850 and the integrity of its associated bus wiring to the IC-600.

(d) Exit the IC-600 self test by pressing the test button on the DC-550.

(e) Using a pitot static tester, simulate 200 kts of airspeed and 1000 ft/minute rate of climb.

(f) Select the IAS mode on the GC-550 and observe the display of approximately 200 kts on thePFD in the air data command location.

(g) Select the VS mode on the GC-550 and observe the display of approximately 1000 ft/minuteon the PFD in the air data command location.

(h) Set simulated airspeed and vertical speed to zero.

7. AH-800 ATTITUDE AND HEADING REFERENCE UNIT

(a) Reference Section I, paragraph 3.3.22, RA Setting 850. AHRS 1 should display VAL DATAindicating a valid AH-800 and its associated wiring to the IC-600

(b) Reference Section I, paragraph 3.3.23, RA Setting 860. AHRS 1 should display VAL DATAindicating a valid AH-800 and its associated wiring to the IC-600

8. PRIMUS II RADIOS

(a) Reference Section I, paragraph 3.3.25, RA Setting 880. VAL DATA should be displayed foreach item, if applicable, indicating a valid radio and the Radio System Bus wiring to theIC-600.



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9. AA-300 RADIO ALTIMETER

(a) Reference Section I, paragraph 3.3.13, RA Setting 760. J2A-049 RAD ALT VALID shoulddisplay 28V.

(b) Reference Section I, paragraph 3.3.7, RA Setting 700. RAD ALT BEL and RAD ALT ABVshould indicate a “V” along with the voltage display.

10. DAU DATA ACQUISITION UNIT

(a) Reference Section I, paragraph 3.3.22, RA Setting 850. J2A-031 R DAU S and J2B-031 LDAU S should both indicate VAL DATA, indicating a valid DAU channel and associated ARINC429 wiring to the IC-600.

(b) Reference Section I, paragraph 3.3.23, RA Setting 860. J2A-041 R DAU P and J2B-081 LDAU P should both indicate VAL DATA, indicating a valid DAU channel and associated ARINC429 wiring to the IC-600.

11. TCAS TRAFFIC COLLISION AVOIDANCE SYSTEM

11.1 PROCEDURES

(a) Ensure that the 115 VAC breaker is in for the TCAS computer.

(b) With the PRIMUS II radio system energized, position the cursor in the TCAS window ofthe RMU.

(c) Press and hold the TST button. The word test will appear in the TCAS window.

(d) Continue holding the TST button. The words TCAS TEST will be displayed indicating thatthe TCAS computer is performing self-test. If the Aural Warning Unit is enabled, a TCAStest advisory will be heard. Additionally, a TCAS Test Pattern will be displayed on the MFD,as shown in Figure 11.1.1.

(e) Continue holding the TST button. Upon completion of the test, the words TCAS PASS ingreen, or ATC ERR in red will appear in the TCAS window. In addition, the TCAS auralwarning will annunciate TCAS PASS or TCAS FAIL separately, as will the MFD annunciateTCAS PASS or TCAS FAIL.



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AD-51728@

NOTE: THE DISPLAY SHOWN MAY NOT REPRESENT ACTUAL FLIGHT CONDITIONS.

N

S

33

15

3012

WE

246

21

3

25

+15+25300

SATTATTAS

FMSKDVT12.512

NMMIN

360

50 50

PLAB1

PLAB2

*PBD01LL01

KDVT

TGTTX-16

STAB TCAS TESTABVFL

MAG2CHK EICAS

RA NO BRGTA NO BRG

31546.0

6

BEZEL MENU DISPLAY AREA

TCASAUTO

+20

-11

+02

-03

Figure 11.1.1. TCAS Test Display



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12. PRIMUS 650/660 WEATHER RADAR

12.1 PROCEDURES

(a) Install a dummy load on the P-650/660 WX Receiver-Transmitter.

(b) Rotate the RADAR knob on the WC-650 WX Controller to TST (Test) position.

NOTE: The radar transmits in all modes except SBY (Standby Mode).

(c) Observe radar test pattern displayed on MFD. It should appear as shown in Figure 12.1.1or Figure 12.1.2. A broken or missing noise band usually indicates a problem with theAFC or magnetron.



360

N

3

6

33

30

5050

FAIL25

TEST MODEFAIL

ANNUNCIATOR

FAULT CODE

TEST PATTERN

25

+15

300

SATTATTAS

FMSKDVT12.512

NMMIN

31547.0

SYSTEM PAGEDISPLAY AREA

AD-51259-R1@

+25

Figure 12.1.1. PRIMUS® 650 MFD Weather Radar Test Mode



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360

N

36

33

30

5050

TEST-10

TEST MODEANNUNCIATOR

TEST PATTERNWITH TEXT

FAULT SHOWN XM

IT O

N! 1F1BB:STRAPS

CO

DE: 27 POC 0 NO STAB SR

CSTA

B UNCAL CHK ATT SR

C

25

+15

300

SATTATTAS

FMSKDVT12.512

NMMIN

31547.0

SYSTEM PAGEDISPLAY AREA

AD-51619@

+25

Figure 12.1.2. PRIMUS® 660 MFD Weather Radar Test Mode



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13. GC-550 FLIGHT GUIDANCE CONTROLLER

(a) Reference Section I, paragraph 3.3.9, RA Setting 720. Activate each pushbutton on theGC-550 controller. Engaging a pushbutton will change the status from ‘0’ to ‘1’. Disengagingthe pushbutton will return the status back to ‘0’. Note that AP and YD are only brought in onthe pilot side IC-600, and therefore will only be displayed if the DU is being driven by thepilot’s IC-600.

(b) Additional GC-550 functions are sent through the DC-550 controller before being received bythe IC-600. See paragraph 4.0

14. PC-400 AUTOPILOT CONTROLLER

These tests verify the Turn Knob and Pitch Wheel functions and their associated wiring betweenthe PC-400 and the IC-600.

(a) Reference Section I, paragraph 3.3.7, RA Setting 700. The TURN KNOB and PITCH WHEELsignals should display a green “V”, valid, and a signal level of approximately 0.00 V dc.

(b) TURN KNOB test: Engage the Autopilot by pressing the AP ENG button on the GC-550.Rotate the Turn Knob fully clockwise. The TURN KNOB readout should be -15 V dc. Rotatethe Turn Knob fully counterclockwise. The TURN KNOB readout should be approximately15 V dc.

(c) PITCH WHEEL test: Engage the Autopilot by pressing the AP ENG button on the GC-550.Rapidly rotate the pitch wheel first in the climb and then in the descend directions. Observe avoltage reading change for the PITCH WHEEL signal.

15. SM-200

During system power up the IC-600 tests the servo motor and tach feedback functions bycommanding the servo to move and then measuring the appropriate servo movement via the tachfeedback signal.

Proper aircraft cable tensions are required for proper autopilot and yaw damper operation. Looseor overly tight cables may result in abnormal autopilot/yaw damper operation.

(a) Engage the autopilot by pressing the AP engage button on the GC-550.

(b) Servo Clutch Test: Reference Section I, paragraph 3.3.19, RA Setting 820. The AP and YDclutch states should change from OPN to HI when the Autopilot is engaged. This checks theIC-600 clutch output status, which is also verified during software power up tests.

(c) With Autopilot engaged ensure servos are engaged by moving the Turn Knob and PitchWheel on the PC-400 and observing the appropriate roll and pitch responses of the aircraftcontrol wheel.



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16. AUTOPILOT AND ELECTRIC TRIM

Pitch trim operation may be compromised by high friction control cables, high breakout forces inthe trim control system, faulty horizontal stabilizer system, or improper trim output signals from theIC-600.

(a) With the autopilot not engaged, manually rotate the aircraft trim wheel to the NOSE DOWNlimit.

(b) Reference Section I, paragraph 3.3.11, RA Setting 740. Move the control wheel trim switchesto the UP position. Pitch trim should rotate 3 complete revolutions in approximately 30seconds and J1A-077 TRIM UP should display 28 V.

(c) Reference Section I, paragraph 3.3.11, RA Setting 740. Move the control wheel trim switchesto the DOWN position. Pitch trim should rotate 3 complete revolutions in the oppositedirection in approximately 30 seconds and J1A-076 TRIM DOWN should display 28 V.

(d) With 0 kts of airspeed, manually rotate aircraft trim wheel to NOSE DOWN limit. Centercontrol column and engage autopilot. Over power the control column in the nose downdirection. After a 1-second delay, the pitch trim wheel should rotate in the NOSE UPdirection, one revolution in approximately 12 seconds. Disengage autopilot.

(e) Simulate greater than 200 kts (190 kts is the switching point for trim speed change). Manuallyrotate the aircraft trim wheel to the NOSE UP limit. Center the control column and engageautopilot. Over power the control column in the nose up direction. After a 1-second delay, thepitch trim wheel should rotate in the NOSE DOWN direction, one revolution in approximately17 seconds.



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(Blank Page)



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SECTION IIIEXPANDED GROUND TEST PROCEDURE

1. INTRODUCTION

The format of the tests is to identify the required response without elaboration. The Work Stepsdefine the preliminary setup for each test but it may be necessary to reset and reengage the modeduring the course of the test. For these reasons these tests should only be used by technicianswho understand the system theory and operation.

All values stated in these tests should have a ± 10% tolerance.

2. PRIMUS 650/660 WEATHER RADAR SYSTEM CHECK

The following paragraphs provide the procedures for operation and checkout of the radar.

2.1 PRELIMINARY CONTROL SETTINGS

Place the system controls on the WC-650 Controller in the following positions before applyingpower from the aircraft electrical system:

MODE OFFGAIN Fully counterclockwiseTILTFully clockwise to +15 positionFunction WX

2.2 PRECAUTIONS

If radar is to be operated in any mode other than STBY , included TEST mode, while the aircraftis on the ground:

(a) Direct the nose of the aircraft so that the antenna scan sector is free of large metallicobjects such as hangars or other aircraft for a distance of 100 feet and tilt antenna fullyupwards.

(b) Avoid operations during refueling of aircraft or other refueling operations within 100 feet.

(c) Avoid operation if personnel are standing too close in the 270 degree forward section of theaircraft. Referring to the Maximum Permissible Exposure Level, this distance would be a5–foot radius from the antenna.

2.3 TEST PATTERN

After warm up, select TEST mode and verify that the test pattern appears as shown in Section IIFigure 12.1.1 or Figure 12.1.2. If the noise band is missing or broken it is indicative of problemswith the system.



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2.4 STABILIZATION INPUT

2.4.1 PRELIMINARY CHECKS

(a) Level the pitch and roll reference axes of the aircraft relative to the earth’s surface.

(b) Verify that the mounting surface of the RTA is aligned to the pitch and roll reference axesof the aircraft within ± ¼ degree.

2.4.2 ELEVATION FEEDBACK CHECK/ADJUSTMENT

(a) Set the XMTR switch on the RTA housing to the OFF position.

(b) On the radar controller, select the STANDBY mode.

(c) Set the SCAN switch on the RTA housing to the OFF position. On the radar controller,adjust the tilt control for 0.0 degree, as displayed on the EFIS/MFD display.

NOTE: Make sure the flat-plate antenna is at 0.0 ± 0.5 degree with respect to theRTA mounting surface.

2.4.3 PITCH GAIN ADJUSTMENT/CHECK

(a) Set SCAN and XMTR switches, on RTA housing, to OFF (toward antenna) and selectStandby on the Controller. The RTA is now in pitch and roll calibration mode.

(b) Select preset GAIN. This forces a 0-degree condition in azimuth by the processor.Azimuth position of the antenna does not matter in calibration mode.

(c) Adjust TILT control on Controller for 0-degree on the face of the flat-plate radiator withan inclinometer.

NOTE: Since the No. 2 AHRS is used to provide the stabilization input to the RTA,press the No. 2 AHRS test button to simulate 10-degree nose up pitch and20 degrees of right roll.

(d) Verify that the antenna tilts down 25 ± 1.5 degrees.

2.4.4 ROLL GAIN ADJUSTMENT/CHECK

(a) Verify that the GAIN control is out of the preset position. This will force a 90-degree rightazimuth condition by the processor.

NOTE: Since the No. 2 AHRS is used to provide the stabilization input to the RTA,press the No. 2 AHRS test button to simulate 10-degree nose up pitch and20 degrees of right roll.

(b) Verify that the antenna tilts up 25 ± 1.5

(c) Return the GAIN control to preset, turn the system off, and set the SCAN and XMTRswitches to ON.

(d) Re-install the vertical reference in its aircraft mounting location.



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2.5 FAULT MONITORS

Critical functions in the RTA are continuously monitored. The RTA utilizes the EFIS as thedisplay mechanism by transmitting detected faults to it via the control bus. The EFIS willtypically display amber WX on the MFD if no SCI data is present or when an RTA failure hasoccurred. This may be due to an actual bus failure or if the RTA is not powered up. When theRTA serial bus is functioning normally, the EFIS will indicate the presence of any failuresdetected by the RTA fault monitoring system and by displaying an amber fault code number inthe tilt angle location of the display when in Test mode.

NOTE: Past faults are stored in the internal memory of the radar. These are accessiblewhen the unit is returned to a Honeywell repair facility. Only the current faults aredisplayed when the unit is placed in Test mode.

The Fault Codes and a brief description of each are listed in Volume I, Section 2.5, Tables 2-5-5and 2-5-6 of this system maintenance manual.

3. EFIS CHECK

The following paragraphs provide the procedures for checkout of the EFIS.

3.1 PRELIMINARY CONTROL SETTINGS

(a) Reversionary Controller: The MFD Selector should be set to NORM. The ADC, SG, andAHRS pushbuttons should not be selected to REV.

(b) Display Controller: BRG 0 and BRG <> knobs selected to OFF.

3.2 DU-870 DISPLAY UNITS

(a) After applying aircraft power, ensure that the displays light up. Use the Dimming Controllerto adjust each display to the desired brightness.

(b) Check each display for color, clarity, and general overall appearance.

(c) Shine a light into each of the photocells on the displays and observe that the displaybrightness level increases.

NOTE: It is normal for the display to overdrive since the photocell is calibrated forchanges in ambient light and not for a direct light source such as a flashlight.



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3.3 DC-550 DISPLAY CONTROLLER

The Display Controller communicates with the Symbol Generator (located in the IC-600) over atwo-wire digital bus. This digital bus can be verified by selecting any function on the controllerand observing the appropriate response on the EFIS displays.

(a) Decision Height (RA) knob: A Decision Height Set Control allows the operator to adjust thedecision height. Rotation of this knob allows the decision height on the PFD to be adjustedbetween 5 and 200 feet in 5-foot increments, and 200 and 990 feet in 10-foot increments.Rotating the RA knob completely counterclockwise removes the decision height from thePFD display.

(b) TEST pushbutton: A system test is provided by a momentary action switch that is actuatedby depressing the decision height (RA) set knob. The TEST pushbutton will place theIC-600 in the test mode. Reference Section 1, paragraph 3.2.2, for expected results whenthe TEST button is activated.

(c) BRG <> and BRG 0 knob test: Rotate the BRG <> and BRG 0 knobs and on the PFD andMFD (Compass and Map modes) observe the bearing pointers and annunciators changesources accordingly:

BRG <>OFFNAV 2ADF (single installation, ADF 2 (dual installation)FMS (option)

BRG 0OFFNAV 1ADF (single Installation, ADF 1 (dual installation)FMS (option)

If the DC-550 is invalid, the onside NAV BRG is displayed by default.

(d) FULL/WX pushbutton: Successive pushes of the FULL/WX button will alternately changethe PFD HSI format from a full compass and ARC mode where radar will be displayed if it isoperating.

(e) GSPD/TTG pushbutton: Actuation of this button will change the display on the PFDbetween GSPD and TTG. If ET is currently displayed, pressing the GSPD/TTG button willselect whichever parameter was previously displayed. Power up state is GSPD.

(f) ET pushbutton: Pushbutton control of an elapsed timer that appears on the PFD in the HSIlocation dedicated to GSPD/TTG. The initial actuation enters the mode at the previoussequence position. Subsequent actuations follow this toggle sequence: RESET -ELAPSED TIME - STOP - REPEAT.

(g) NAV pushbutton: This pushbutton allows selection and toggling of short range navigationsources. The sources to be toggled are: Pri VORLOC - Sec VORLOC.

(h) FMS pushbutton: Allows selection of LRN as the NAV source.



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3.4 DIMMING CONTROLLERS

The Dimming Controllers consist of five rotary knobs which interface with each of the DUs suchthat they can be independently dimmed. Verify that each knob allows brightness control of itsrespective DU, and that with all the knobs in the minimum brightness position that all the DUsappear of equal intensity.

3.5 REVERSIONARY PANEL

(a) Press the SG REV pushbutton on the Pilot’s Reversionary Panel and verify that an amber“SG2” is displayed in the upper left-hand corner of each PFD.

(b) Press the SG REV pushbutton on the Copilot’s Reversionary Panel and verify that an amber“SG1” is displayed in the upper left-hand corner of each PFD.

(c) Press the ADC REV pushbutton on the Pilot’s Reversionary Panel and verify that an amber“ADC2” is displayed in the upper left-hand corner of each PFD.

(d) Press the ADC REV pushbutton on the Copilot’s Reversionary Panel and verify that anamber “ADC1” is displayed in the upper left-hand corner of each PFD.

(e) Press the AHRS REV pushbutton on the Pilot’s Reversionary Panel and verify that anamber “ATT2” is displayed in the middle left-hand corner of each PFD.

(f) Press the AHRS REV pushbutton on the Copilot’s Reversionary Panel and verify that anamber “ATT1” is displayed in the middle left-hand corner of each PFD.

3.6 MFD BEZEL BUTTONS AND ROTARY SET KNOB

Six bezel buttons are located on the bottom of the MFD display along with a rotary set knob.The menu for the bezel buttons is shown above the buttons on the bottom of the MFD CRT.These menus will change depending on the mode that is selected on the MFD. The bezel buttontree menu is shown in Figure 3.6.1.

(a) Main Bezel Button Menu: The main bezel button menu is the default power up.

(b) Press the SYS button and verify that the menu changes to the SYS Submenu. Press theRTN button and verify that the menu returns to the main bezel menu. Repeat thisprocedure for each of the remaining buttons on the main bezel button menu.

(c) With the main bezel button menu displayed, press the MAP/PLAN button and then rotatethe M/P RNG knob and observe the range change on the MFD.



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A C B

AD-51208@

SYS MFD CKLSTM/PRNGTCAS WX

MAPPLAN

MENU

RTN SKP LNBKM/PRNGPAG RCL ENT

CHECK LIST

RTN SKP RCL ENT

JOYSTICK

RTN T/OECSA/I

M/PRNGFUEL HYD ELEC

RTNNAVAPT DATA

M/PRNGSPDS MAINT

MENU

RTNNAVAPT DATA

M/PRNGSPDS JSTK MAINT

MENU

RTNV189 SET

VR---

V2---

AP---

M/PRNG

Figure 3.6.1. MFD Bezel Button Menu Tree



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4. MADC CHECK

The MADC has two operating modes: normal and self-test. The normal mode is for normalaircraft operation. The self-test mode is initiated by applying a DC ground to the self-test pin onthe rear connector. This would normally be interlocked with a WOW switch, and is inhibited whenmore than 50 knots of airspeed is present. In the self-test mode, the outputs of the MADC aredriven to preset values to check the operation of the MADC, interconnects, and the indicators.

4.1 NORMAL MODE

(a) After applying power to the system, verify on the PFD that the Airspeed, Altitude, andVertical Speed displays show valid information from the MADCs.

(b) Press the ADC REV pushbutton on the Pilot’s Reversionary Panel and verify that the Pilot’sPFD shows “ADC 2” in the upper left-hand corner and that the Airspeed, Altitude, andVertical Speed displays show valid information.

(c) Press the ADC REV pushbutton on the Copilot’s Reversionary Panel and verify that theCopilot’s PFD shows “ADC 1” in the upper left-hand corner and that the Airspeed, Altitude,and Vertical Speed displays show valid information.

(d) Press the ADC REV pushbuttons again to return the displays to their normal configuration.

(e) Press the IN/HPA pushbutton on the PFD bezel controllers and verify that the BarometricCorrection display on the respective PFDs toggle between IN (inHg) or HPA(HectoPascals).

(f) Rotate the BARO knobs on the PFD controllers and verify that the Barometric CorrectionDisplays on the respective PFDs change value either by 0.01 inHg per click or 1 HPA perclick, depending on which is selected.

(g) Press the STD pushbutton on the PFD bezel controllers and verify that the BarometricCorrection Displays on the respective PFDs display standard baro correction values of29.92 in inHg is selected or 1013 HPA if HectoPascals is selected.

(h) Using a pitot static tester, simulate 200 knots of airspeed and 1000 ft/min rate of climb.Verify that the Airspeed and Vertical Speed Displays on the PFD show the simulated valuesfrom the pitot static tester. Press the GC-550 CPL pushbutton to select the side to betested (left or right). The MADC selected for display on the side selected will be the MADCtested in steps (i) and (j)

(i) Press the SPD pushbutton on the GC-550 Flight Guidance Panel and verify that the currentairspeed is displayed as the Airspeed Target above the Airspeed Tape on the PFD (foraltitudes less than 34,000 feet). The target switches automatically from indicated airspeedtarget to mach number target as the system climbs through 34,000 feet. The targetautomatically switches from mach number target to indicated airspeed target as the aircraftdescends through 33,750 feet.

(j) Press the VS pushbutton on the GC-550 and verify that the Vertical Speed CommandDisplay and Command Bug on the Vertical Speed Display on the PFD displays the currentvertical speed.

(k) Return the simulated airspeed and vertical speed to zero.

(l) Press the CPL pushbutton on the GC-550 to indicate a left arrow. This will select the leftMADC as the master for altitude preselect operation. Also verify that MADC 1 is selectedfor display on the left PFD and MADC 2 is selected for display on the right PFD.



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(m) Rotate the ASEL knob on the GC-550 and verify that the Altitude Preselect Display value onboth PFDs changes correspondingly.

(n) Press the CPL pushbutton on the GC-550 to indicate a right arrow. This will select the rightMADC as the master for altitude preselect operation. Also verify that MADC 1 is selectedfor display on the left PFD and MADC 2 is selected for display on the right PFD.

(o) Rotate the ASEL knob on the GC-550 and verify that the Altitude Preselect Display value onboth PFDs changes corresponding.

4.2 SELF-TEST MODE

(a) Place system in WOW mode (Weight-On-Wheels) and apply a DC ground to MADC pinJ1-70 (Functional Test Discrete). Ensure that airspeed is less than 50 knots. This placesthe MADC in Self-Test mode.

(b) The MADC being tested will output the following on the respective PFD and MFD.

OUTPUTS DATA VALUE

Pressure Altitude 4000 Feet

Baro-Corrected Altitude 1000 Feet

Altitude Rate 5000 Ft./Min

Calibrated Airspeed (CAS) 325 knots

True Airspeed (TAS) 325 knots

Mach 0.77 M

Static Air Temperature (SAT) -45 deg C

Total Air Temperature (TAT) -16 deg C

Baro-Correction inHg 29.92 inHg

Baro Correction hPa 1013 hPa

Static Pressure 29.92 inHg

Total Pressure 29.92 inHg

Dynamic Pressure 0 inHg

Max Operating Speed (VMO) 0.76 M

Vmo Warning Active (GND) for 2 sec

MADC Valid Inactive (OPEN)

Programmable Output Discretes Active (GND)



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5. RADIO ALTIMETER TEST

CAUTION: UNDER NO CIRCUMSTANCES SHOULD POWER BE TURNED ON WITH THETRANSMIT ANTENNA DISCONNECTED FROM THE TRANSMITTER.

The following paragraph provide the procedures for the checkout of the Radio Altimeter System.

Step 1: Depress and hold the TEST button on the DC-550 Display Controller to place the RadioAltimeter system in TEST.

Step 2: Set the Decision Height display on the PFD to 0 feet, using the Decision Height (RA) knoband place the system in the “air” mode.

Step 3: Press and hold the TEST button for a minimum of 5 seconds. The Radio Altimeterdisplay on the PFD should increase to 100 ± 10 feet.

Step 4: Observe that a box is drawn.

Step 5: Release the Test pushbutton and observe that “MIN” is drawn inside the box and flashesfor 10 seconds.

6. AHRS CHECK

The AHRS has six fundamental operating modes. They are Initialization, Full Performance, DG,Basic, Test, and Maintenance modes.

The initialization mode is entered upon power up of the system. During this mode the systemperforms self tests to determine the condition of its components (sensors, CPU, power supply,etc). At the end of the initialization process, the system enters full performance mode unless asystem input (lack of input) or operator command has placed it in one of the reversionary modes(basic or DG).

The test mode can be activated by the operator at any time while the aircraft is on the ground.During the test the system outputs are driven to preset test values to verify proper operation of thedata channels, interconnects, and indicators.

The maintenance mode is designed for use in installing and maintaining the AHRS. An installeruses the mode for setting discretes (orientation, SDI, and ARINC 29 Hi/Low Select), determiningthe flux valve compensation coefficients, and determining the mounting tray alignment coefficients.Honeywell field support uses the mode for special data retrieval on the aircraft. A special functionin the maintenance mode is the automatic compass swing. This mode is accessed by groundingthe Maintenance Test pin on the AHRU and communicating via RS-232 with a laptop andHoneywell supplied software.



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6.1 AHRS POWER-UP CHECKOUT

Verify the AHRS Power-up by performing the following:

(a) Make sure that all AHRS circuit breakers are set to OFF.

(b) Set the AHRS 28 V dc circuit breaker to ON.

(c) Verify that the AHRS cooling air fan has turned on by checking for airflow at the air exitholes in the AHRS cover.

(d) Verify that the PFD is displaying valid Attitude and Heading information. Allow 2 minutes forthe AHRS to become valid.

6.2 AHRS TEST MODE

(a) Place the system in WOW (Weight-On-Wheels) mode.

(b) Place the AHRS in Test mode by pressing the remote mounted AHRS TEST pushbutton onthe maintenance panel.

(c) Verify the AHRS outputs the correct test values on the respective PFD and MFD.

6.3 AHRS ATTITUDE OUTPUTS CHECK

Verify the AHRS attitude by performing the following:

(a) Place the aircraft on jacks and level the aircraft per the manufacturer’s applicable aircraftleveling procedure. The aircraft should be leveled to within ± 0.1 degree.

(b) Raise or lower the nose of the aircraft until it is 2 degrees away from the level position.

(c) Verify that the ADIs on the PFDs display a minimal change in roll angle.

(d) Verify that the ADIs on the PFDs indicate a pitch with the correct polarity.

6.4 AHRS FLUX VALVE CHECK

Verify the AHRS flux valve by performing the following:

(a) Move the aircraft to a level site that is free of buried metal objects and is at least 200 feetfrom the nearest metal structure.

(b) Move the aircraft around a 360-degree compass swing. During the compass swing verifythat the heading output of the AHRS on the HSI of the PFDs correctly indicates the headingof the aircraft to within ± 1 degree throughout the compass swing.



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6.5 VERIFICATION OF MEMORY MODULE PROGRAMMING

If the previous AHRS CHECK procedures have been performed then it has been verified that thememory module on the mounting tray has been programmed correctly.

The programming of each location in the memory module was verified during the AHRScheckout procedure as described below.

(a) Memory Module Programmed Flag. The AHRS will not enter the flux valve calibrationprocedure unless the Memory Module Programmed Flag in the memory module has beenprogrammed correctly. If the flux valve calibration was completed successfully then it hasbeen verified that the Memory Module Programmed Flag in the memory module has beenprogrammed correctly.

(b) Orientation Select. If this location is not programmed correctly, the AHRS pitch and rolloutputs will be incorrect. If the orientation stored in the memory module is 90 degrees offfrom the actual orientation of the AHRS, then the pitch of the aircraft will appear in theAHRS roll output, and the roll of the aircraft will appear in the AHRS pitch output. If theorientation stored in the memory module is 180 degrees off from the actual orientation ofthe AHRS, then the pitch and roll outputs of the AHRS will have the wrong sign. (i.e., theywill be positive when they should be negative and vice versa.) If the AHRS Attitude OutputsCheckout procedure described in Section 6.3 was completed successfully, then theorientation select in the memory module has been programmed correctly.

(c) Source/Destination Identifier. If this location is not programmed correctly, then the AHRSwill not respond when it is addressed on the ASCB or ARINC bus. If the AHRS ARINCCheckout (Section I paragraph 3.3.22 and 3.3.23, RA 850 and 860) procedure wascompleted successfully, then the source/destination identifier in the memory module hasbeen programmed correctly.

(d) ARINC Configuration. If this location is not programmed correctly, then the AHRS will notrespond when it is addressed on the ARINC bus. If the AHRS ARINC Checkout (Section Iparagraph 3.3.22 and 3.3.23, RA 850 and 860) procedure was completed successfully, thenthe ARINC configuration in the memory module has been programmed correctly.

(e) Controller Configuration. If this location is not programmed correctly, then the AHRS willnot respond to controller inputs. If the AHRS was successfully placed in the output testmode via a controller command during the AHRS Test Mode (Section 6.2) procedures thenthe controller configuration in the memory module has been programmed correctly.

(f) Mounting Tray Misalignment Angle. If these locations are not programmed correctly, theAHRS pitch and roll outputs will be incorrect. If the AHRS Attitude Outputs Checkoutprocedure described in Section 6.3 was completed successfully, then the mounting traymisalignment angle in the memory module has been programmed correctly.

(g) Discrete Data Sumcheck. If this location is not programmed correctly the memory modulesumcheck power-on BITE test in the AHRS will fail, and the AHRS outputs shall indicate anAHRS fault. If the AHRS was successfully powered up without the outputs indicating anAHRS fault, then the discrete data sumcheck in the memory module is correct.



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(h) Flux Valve Calibration Coefficients. If these locations are not programmed correctly, theoutput of the AHRS flux valve will not be properly compensated for errors caused bysources of magnetism other than the earth’s magnetic field. If the AHRS Flux ValveCheckout (Section 6.4) procedure was completed successfully, then the flux valvecalibration coefficients must have been correctly stored in the memory module.

(i) Flux Valve Calibration Date. These locations are not used by the AHRS, therefore there willbe no faults or errors if these locations are not programmed correctly, as long as the fluxvalve calibration data sumcheck in the memory module accurately reflects the flux valvecalibration date data.

(j) Flux Valve Calibration Coefficients Sumcheck. If this location is not programmed correctlythe memory module sumcheck power-on BITE test in the AHRS will fail, and the AHRSoutputs shall indicate an AHRS fault. If the AHRS was successfully powered up without theoutputs indicating an AHRS fault, then the flux valve calibration coefficients sumcheck inthe memory module is correct.



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7. PRIMUS II RADIOS CHECK

This section provides procedures for checking the PRIMUS® II, SRZ-85X Series Integrated RadioSystem for the correct installation and proper operation of all components in the system.

7.1 BUS OPERATION CHECKOUT

Step 1: With all power off, verify that the remote radio units are installed in their respectivemounting trays. Install all other radio units and connect them to their respectivemating connector(s).

Step 2: Apply 28 V dc power to System 1 RMU.

Step 3: Verify that upon initial application of power, the RMU page comes up forapproximately 8 seconds during which time the Power On Self Tests (POST) arebeing performed.

Step 4: If no errors are detected, the main tuning page will be displayed after completion ofthe POST. If an error is detected, it will be displayed on either an RMU/SYSTEMTEST RESULTS page or a RADIO TEST RESULTS page. If an error is displayed,ignore the test results at this time and go to the main page by momentarilypressing the TST button on the RMU.

Step 5: Select the Maintenance Menu page on the RMU by simultaneously pressing theCOM and NAV transfer buttons (top button on each side of the RMU).

Step 6: Select the RMU SETUP menu. The RMU may indicate “MLS ENABLED.” TheML-850 MLS Receiver is not part of the EMB-145 radio system, thereforemomentarily press the line select button adjacent to the “MLS ENABLED” display.The display should change to “MLS DISABLED.”

Step 7: Return to the main radio page as indicated by the instructions on the RMU. If theMLS window was turned off, the MLS window on the RMU should be blank. If theMLS window was left on, the MLS display window should be on.

Step 8: Apply 28 V dc power to System 1 Com Units by closing the COM 1 and ATC 1circuit breakers.

Step 9: Momentarily press the COM line select button on the RMU and verify that theyellow cursor box moves to line 2 of the COM window.

Step 10: Rotate the RMU tuning knobs and verify that the COM frequency does change.Press the COM transfer button and verify that the active and the temporaryfrequencies switch locations.

Step 11: Press and hold the COM line select button on the RMU and verify that the yellowcursor box moves to line 1 of the COM window for direct tuning. Press and holdthe line select button to return to normal tuning.

Step 12: Momentarily press the transponder code line select button on the RMU and verifythat the yellow cursor box moves to line 1 of the ATC window. Cursor automaticallyreturns to the COM window after 20 seconds of inactivity.

Step 13: Rotate the RMU tuning knobs and verify that the ATC codes change.

Step 14: Press and hold the transponder line select button on the RMU and verify that theactive ATC code is replaced by the stored ATC code.



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Step 15: Momentarily press the transponder mode line select button on the RMU and verifythat the yellow cursor box moves to line 2 of the ATC window.

Step 16: To change the memory, change the code using the RMU tuning knobs, press store(STO),change the code to another value, press and hold the transponder modeline select button again and the stored value should be should be displayed.

Step 17: Rotate the RMU tuning knobs and verify that the ATC operating modes change. Ifthe transponder is in STBY mode, rotating the tuning knobs will have no effect.

Step 18: Momentarily press the transponder mode line select button on the RMU and verifythat the mode changes from the previously selected System 1 mode (ATC ON,ATC ALT, TA/RA, TA ONLY) to STANDBY.

Step 19: Apply 28 V dc power to System 1 NAV Unit by closing the VOR/ILS 1, DME 1, andADF 1 circuit breakers.

Step 20: Momentarily press the NAV line select button on the RMU and verify that theyellow cursor box moves to line 2 of the NAV window. Cursor automatically returnsto the COM window after 20 seconds of inactivity.

Step 21: Rotate the RMU tuning knobs and verify that the NAV frequency does change.Press the NAV transfer button on the RMU and verify that the active and temporaryfrequencies switch locations.

Step 22: Press and hold the NAV line select button on the RMU and verify that the yellowcursor box moves to line 1 of the NAV window for direct tuning. Press and hold theline select button to return to normal tuning.

Step 23: Momentarily press the DME button on the RMU and verify that the NAV windowgoes into a split mode with DME on the bottom (line 2)

Step 24: Rotate the RMU tuning knobs and verify that the DME frequency changes.

Step 25: Momentarily press the DME button on the RMU again and verify that the DMEfrequency switches to a Tacan channel number.

Step 26: Rotate the RMU tuning knobs and verify that the Tacan channel numbers change.

Step 27: Momentarily press the DME button on the RMU again and verify that the NAVwindow is no longer in the split mode.

Step 28: Momentarily press the ADF frequency line select button on the RMU and verify thatthe yellow cursor box moves to line 1 of the ADF window. Press and hold the lineselect button to view the memory value.

Step 29: Rotate the RMU tuning knobs and verify that the ADF frequency changes.

Step 30: Momentarily press the ADF frequency mode select button on the RMU and verifythat the yellow cursor box moves to line 2 of the ADF window. Cursor automaticallyreturns to the COM window after 20 seconds of inactivity.



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Step 31: Rotate the RMU tuning knobs and verify that the various ADF operating modes(ADF, BFO, VOICE, ANT) cycle through the window.

Step 32: Momentarily press the ADF mode line select button on the RMU and verify that themode changes from the previously selected System 1 mode (ADF, BFO, VOICE,ANT) cycle through the window. Cursor automatically returns to the COM windowafter 20 seconds of inactivity.

Step 33: If any of the frequency, code, or mode functions fail (dashes in cursor box), themost probable cause will be a miswired bus. Remove all power and check buswiring.

Step 34: Select the Maintenance Menu page on the RMU by simultaneously pressing theCOM and NAV transfer buttons (top button on each side of the RMU). Verify thatthe RMU leaves its normal operating mode and displays a menu allowing theoperation to enter the Aircraft Maintenance Mode.

Step 35: Select the STRAP menu and check the strap display of each unit (COM UNIT,NAV UNIT, and RMU) to ensure that the proper strap options have beenincorporated per the installation diagrams.

Step 36: Repeat steps (a) through (hh) for System 2 radio components.

Step 37: Remove all radio system power.

7.2 UNIT POWER ON SELF TEST (POST)

Step 1: Ensure that all radio system power is OFF.

Step 2: To ensure proper POST sequencing, apply 28 V dc power to all System 1components, ensuring that the power is applied to all components simultaneously.

Step 3: Upon initial application of power to the RMU, the system will automatically performa Power On Self Test (POST). If no errors are detected, the main tuning page willthen be displayed. If an error is detected, it will be displayed on either anRMU/SYSTEM TEST RESULTS page or a RADIO TEST RESULTS page.

Step 4: Annotate all errors for later troubleshooting. Momentarily press the TST button onthe RMU to proceed to the main tuning page.

Step 5: Repeat steps (a) through (d) for System 2 radio components.



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7.3 AIRCRAFT INTERFACE/RAMP TESTS

7.3.1 COM 1 TESTS

Step 1: Ensure that the RMU cursor box is in the COM window. Press and hold the TSTbutton on the RMU and verify that COM self test is being performed. If the COMself test does not pass, annotate failure for later troubleshooting.

Step 2: Tune COM 1 to any COM frequency and momentarily press the SQ button onthe RMU. Verify that SQ is displayed in the banner of the RMU COM windowand that noise can be heard in the audio system.

Step 3: Momentarily press the SQ button on the RMU again and verify that audio noisedisappears.

Step 4: Using the RMU tune COM 1 and 2 to a local ground station or to a hand heldCOM. Choose the frequency to avoid blocking a busy channel.

Step 5: On the System 1 audio panel, select the COM 1 microphone switch. Set theCOM 1 volume control to mid range.

Step 6: Momentarily key a System 1 microphone and verify that a TX is displayed in thebanner of the RMU COM 1 window.

Step 7: Transmit to the local ground station or hand held requesting a radio check.

Step 8: Verify proper transmission and reception of audio.

Step 9: Verify sidetone is working (hear yourself talk in headphones and optionally inspeaker).

Step 10: Select the COM 2 MIC select switch on the System 1 audio panel.

Step 11: Repeat steps (f) through (i).

Step 12: Select the EMER switch on the System 1 audio panel.

Step 13: Repeat Steps (f) through (I).

Step 14: Repeat COM Tests for System 2 equipment if applicable.



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7.3.2 TRANSPONDER 1 TESTS

Step 1: Ensure that the RMU cursor box is in the ATC window. Press and hold the TSTbutton on the RMU and verify that the ATC self test is being performed. If theATC self test does not pass, annotate failure for later troubleshooting.

Step 2: Using the RMU, set the transponder code to a local VFR and mode to 1 ON.

NOTE: When selecting a transponder code, be careful not to radiate any of thefollowing codes unintentionally.

Code 0001 (Military intercept code)Code 7500 (Hijack code)Code 7600 (VHF COM receiver failure code)Code 7700 (Emergency code)

Step 3: Using an ATC ramp tester, verify that the local VFR code is being transmitted.

Step 4: Verify that the reply annunciator, located in the RMU upper right corner of theATC window, comes on.

Step 5: Momentarily press the ID button on the RMU and verify that ID is displayed inthe banner of the RMU ATC window. Verify that the ATC ramp tester ID lightcomes on.

Step 6: Using the RMU, set the transponder mode to 1 ATC ALT.

Step 7: Set the aircraft’s cockpit barometric altimeter to indicate 29.92 inHg (1013 mb)and verify that the altitude code segment of the ATC ramp tester indicates theapproximate field elevation.

Step 8: Verify that the Transponder automatic reporting output displayed on the ramptester and the altitude displayed on the barometric altimeter does not exceed125 feet.

Step 9: Check transponder sensitivity by placing the ramp tester antenna approximately10 feet from the aircraft transponder antenna. Verify that reply rate is graterthan 90 percent.

NOTE: Step 10 needs to be accomplished only at initial transponder installation and aspart of the FAA mandated 24 calendar month static system maintenancerequirements.

Step 10: Using an air data test fixture, slowly run up the altitude to the aircraft’smaximum allowable ceiling. Verify that the baro altimeter indicator indicates thesame altitude as the ATC ramp tester.

Step 11: Repeat Transponder Tests for System 2 equipment if applicable.



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7.3.3 NAV 1 TESTS

Place the RMU cursor box in the NAV window. Press and hold the TST button on the RMUand verify that the VOR/DME self test is being performed. Also verify that the navigationdisplays indicate as follows:

Marker Marker Beacon annunciators flash on in the following sequence:Outer Marker- BlueMiddle Marker- AmberInner Marker- White

LOC/GS Glide Slope deviation needle/bar moves up 1 standard deviation (dot).

VOR With the OBS set to zero degree, the course deviation bar should centerand the TO/FROM annunciator should indicate TO.

DME DME displays 10.0 NMI.

If the VOR/DME self test does not pass, annotate the failure for later troubleshooting.

7.3.3.1 VOR Tests

Step 1: Using a VOR ramp tester, set up a frequency of 108.00 Mhz with a bearing of0 degree TO the station.

Step 2: Using the RMU, tune the NAV to 108.00 Mhz. Verify that the aircraft bearingindicator indicates 0 degree TO the station with the VOR Flag out of view.

Step 3: Rotate OBS 10 degrees and verify full scale deflection of left/right needle.

Step 4: Switch ramp test TO-FROM switch to FROM. Verify that the bearingindicator TO-FROM annunciator displays 0 degree FROM the station.

Step 5: Set up ramp tester to transmit an audio Ident signal. Verify presence of identtone on aircraft audio system. (Verify ident filter is off on Audio System.)

Step 6: Repeat above steps using the 90, 180, and 270-degree radials.



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7.3.3.2 LOC 1 Tests

Step 1: Using a LOC ramp tester, set up to transmit a 108.10 Mhz on course signal.

Step 2: Using the RMU, tune the NAV to 108.10 Mhz. Verify that the aircraft localizerdeviation indicator indicates on course with its flag out of view.

Step 3: Set up LOC ramp tester to transmit left (0.093 DDM) of course deviation.Verify that the aircraft deviation indicator indicates left of course (One dot).

Step 4: Set up LOC ramp tester to transmit a right (0.093 DDM) of course deviationsignal. Verify that the aircraft deviation indicator indicates right of course(One dot).

Step 5: Set up ramp tester to transmit an audio signal. Verify the presence of identtone on aircraft audio system.

Step 6: Set up ramp tester to transmit 108.00 Mhz. Verify that the aircraft LOCindicator Flag comes into view.

7.3.3.3 Glide Slope 1 Tests

Step 1: Using a Glide Slope ramp tester, set up to transmit a 108.10 Mhz On Coursesignal.

Step 2: Using the RMU, tune the NAV to 108.10 Mhz. Verify that the aircraft GSdeviation indicator indicates on course with its flag out of view.

Step 3: Adjust Ramp Generator Modulation for 0.091 DDM (Up needle). Verify thataircraft GS course needle is a half scale (one dot) up.

Step 4: Adjust Ramp Generator Modulation for 0.091 DDM (Down needle). Verifythat aircraft GS course needle is at half scale (one dot) down.

Step 5: Set up ramp tester to transmit a frequency of 108.00. Verify that the GSindicator flag comes into view.

7.3.3.4 Marker Beacon 1 Tests

Step 1: Using a Marker Beacon ramp tester, switch on the outer marker. Verify thatthe aircraft’s blue marker light is activated and that a 400 Hz audio tone ispresent on the aircraft audio system.

Step 2: Switch on the ramp tester middle marker. Verify that the aircraft’s ambermarker light is activated and that a 1300 Hz audio tone is present on theaircraft audio system.

Step 3: Switch on the ramp tester inner marker. Verify that the aircraft’s whitemarker light is activated and that a 3000 Hz audio tone is present on theaircraft audio system.

Step 4: Repeat NAV Tests for System 2 NAV equipment if applicable.



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7.3.4 DME 1 TESTS

Step 1: Using a DME ramp test, set up to transmit on 108.00 Mhz and set in 35 milesand 200 knots to the station.

Step 2: Using the RMU, select DME split mode (DME Hold) and tune DME to afrequency of 108.00 Mhz. Verify on the aircraft’s DME display a distancedecreasing from 35 miles. Verify a DME Hold annunciation.

Step 3: Activate ident tone on ramp tester. Verify tone on aircraft audio system on boththe headphone and speaker outputs.

Step 4: Repeat DME Tests for System 2 equipment if applicable.

7.3.5 ADF 1 TESTS

NOTE: Reliable ADF reception and bearing indications cannot be assured unless theaircraft is removed from noise sources, such as power carts and fluorescentlighting. The aircraft should also be removed from large metallic structuressuch as hangars and power lines that may distort or attenuate the radio fields.

Step 1: Using the RMU, tune in a local non-directional beacon or AM radio stationfrequency if a beacon is unavailable using the ADF frequency window.

Step 2: Select BFO mode in the RMU ADF mode select window. Verify the reception ofan audio signal with a beat frequency on the aircraft audio system. Verify thatthe aircraft’s ADF bearing pointer points to the station.

Step 3: Select ADF mode in the RMU ADF mode select window. Verify the reception ofan audio signal on the aircraft audio system. The ADF needle should point tostation.

Step 4: Select ANT mode in the RMU ADF mode select window. Verify the reception ofan audio signal on the aircraft audio system. Verify that the aircraft ADFbearing pointer is removed from the EFIS display.

Step 5: Tune ADF to a strong local broadcast station frequency. Select VOICE modeand verify extended audio frequency range.

Step 6: Repeat ADF Tests for System 2 equipment if applicable.



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7.3.6 RMU TESTS

7.3.6.1 Dimming Check

Step 1: Momentarily press the DIM button on the RMU and verify that the dimmingpattern is displayed. Rotate the tuning knobs and verify that dimming occurs.

Step 2: Adjust the dimming either up or down to suit operator’s preference. To leavethe DIM mode, press any line select button or momentarily press the DIMbutton again.

7.3.6.2 Power ON/OFF Check

Step 1: Pull the circuit breakers for each radio unit. Return to the main tuning pageand verify that the radios frequencies, codes, and channel windows displaydashes indicating that power to the remote unit is off.

Step 2: Return power to all units. Return to the main tuning page and verify that allfrequencies, codes, and channels are displayed as before.

NOTE 1: Repeat RMU Tests for System 2 equipment if applicable.

NOTE 2: If dual radio systems are installed, momentarily press the 1/2 button on theRMU. Check cross-side operation by verifying that RMU No. 1 (pilot’s) cancontrol System No. 2 (copilot’s) radios.

NOTE 3: Repeat above step by verifying that RMU No. 2 (copilot’s) can control SystemNo. 1 (pilot’s) radios.



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7.3.7 TUNING BACKUP CONTROL HEAD (CLEARANCE DELIVERY HEAD - CDH) TESTS

NOTE: The following tests apply only if the CD-850 Clearance Delivery CDU isconfiguration strapped for clearance delivery mode.

Step 1: Check Clearance Delivery CDU display for any Power On Self Test (POST)error codes. Annotate any failure codes and then press any switch or rotate anyknob to exit POST mode.

Step 2: Set the Clearance Delivery CDU mode select switch to normal. The COM andNAV frequencies displayed on the Clearance Delivery CDU should be the sameas the active frequencies shown in the RMU COM and NAV windows.

Step 3: Set the Clearance Delivery CDU mode select switch to emergency. Verify thatEMRG is displayed vertically in the upper-right corner of the CDU display.Verify that AUX ON is displayed in the RMU COM and NAV windows.

Step 4: Press the SQ button on the Clearance Delivery CDU. Verify that the SQannunciator on the CDU display turns on and that noise can be heard in theaudio system. Press SQ button again; the SQ annunciator should turn off.

Step 5: Tune the Clearance Delivery CDU to a local VHF communications frequencyand verify clear and understandable reception and transmission.

Step 6: Tune the Clearance Delivery CDU to a local VOR station and verify that theproper VOR signals are being received.

Step 7: Press the NAV AUDIO button on the Clearance Delivery CDU. Verify that theNAV AUDIO annunciator on the CDU display turns on and that normal receiveraudio can be heard in the audio system.

Step 8: Press the NAV AUDIO button again; the NAV AUDIO annunciator should turnoff.

Step 9: Set the Clearance Delivery CDU mode select switch to normal. The EMRGannunciator should turn off.

7.4 SYSTEM STATUS MESSAGE CHECKS

The procedures in this section are to be used to check the maintenance log to determine if anyerrors have been entered into the log.

Step 1: Simultaneously press the COM and NAV transfer buttons (top button on each sideof RMU). Verify that the RMU display leaves its normal operating mode anddisplays a menu which allows the operator to enter the Aircraft Maintenance Mode(AMM).

Step 2: Select the MAINTENANCE LOG menu and check the log of each component forany errors that may have been entered during the Power On Self Test (POST) andPilot Activated Self Test (POST) performed in Sections 7.2 and 7.3. Refer toSection 5 of the PRIMUS II Maintenance Manual, Pub. No. A15-3800-01, FaultIsolation, for troubleshooting procedures.



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7.5 IN-FLIGHT TESTS

The in-flight test procedures are not mandatory and it is recommended that they be performedonly to check for a suspected fault that cannot be checked on the ground. If in-flight checks arenecessary they can be performed in any sequence desired. In addition only selected tests needbe performed.

7.5.1 AUDIO PANEL FLIGHT TEST PROCEDURES

• Electromagnetic Compatibility - With all systems operating in-flight, verify by observation,that operation of the Audio Panel has no adverse effects on other aircraft systems.

• Audio Quality Check - Monitor all audio signals during flight, testing all individual radiosincluding but not limited to:

− Intelligibility of VHF communications

− Localizer ident

− Marker beacon audio

− VOR ident

− ADF ident

− DME ident

• Emergency Com Check - Depress “Emer Com” button and verify operation of audio panelin a simulated failure mode.



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8. TRAFFIC COLLISION AVOIDANCE SYSTEM (TCAS)

8.1 TCAS CHECK

The TCAS II is programmed to perform automatic self-test at power-on and during operation.Resident test programs are executed independently in the computer unit. The test performedinclude power-on self-test, continuous self-test, and commanded self-test.

The computer unit detests system faults and reports them on its front panel lamp display. Itsflight leg memory stores status and fault information for 10 consecutive flight legs. A flight leg isthe interval between weight-off-wheels and weight-on-wheels during which the TCAS isoperative. By recalling the stored data, ground maintenance personnel can evaluate in-flightperformance on the ground and fault isolate a current or previous failure to a specific LRU orLRU interface.

Table 8.1 lists the functions of the computer unit display lamps and the correspondingtroubleshooting actions. If the lamps indicate an antenna problem, the antenna connectionsshould be checked by measuring wiring resistance at the tray per Figure 8.1.1 and Table 8.2.The resistance values in Table 8.2 are between center conductor and shield on each LTP andLMP connector.

Table 8.3 summarizes how computer unit self-test is executed at power-on, during TCASoperation, and during commanded self-test. The computer unit can execute commandedself-test only when the aircraft is on the ground.



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PWR/GNDSIGNAL

LTP

LMP

LBP

12

34RTP

RMP

RBP

BLANK

SIGNAL

AD-45492@

12

34

10

OHM METERLEADS

Figure 8.1.1. Antenna Test Points (Rack Side)



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Table 8.1. CU Fault Reporting and Corrective Actions

LAMP FAILURE IDENTIFICATION POSSIBLE ACTION

TCASPASS

The TCAS II was operational inthe flight displayed.

N/A

TCASFAIL

TCAS failed in the flightdisplayed.

Remove and replace TCAS CU, if there is noother FAIL indication.

TOP ANT Top antenna, or cablingconnectors to top antenna arefailed or miswired.

Verify antenna cabling and connections byverifying resistance at TCAS CU tray is perFigure 8.1.1 and Table 8.2. Antenna WiringResistance(after removing TCAS CU). Repaircabling, or replace antenna, as required.

BOT ANT Bottom antenna, or cabling andconnectors to bottom antenna,are failed or miswired.

Verify antenna cabling and connections byverifying resistance at TCAS CU tray is perFigure 8.1.1 and Table 8.2. Antenna WiringResistance(after removing TCAS CU). Repaircabling, or replace antenna, as required.

HDG Heading input function is notused.

N/A

RA LOG Function is not used. N/A

TA DISP The LRU providing the trafficadvisory display function (on oneor both sides) is not available toTCAS.

Verify wiring and power to TA display (on bothsides of the cockpit, if two are installed). TAdisplay may be provided on VSI/TRA, WX radarindicator, or EFIS. Check that TA validdiscretes 1 and 2 (RMP-7E and RMP-7J) are 10ohms to ground maximum. Repair wiring, orreplace TA display, as required.

RA DISP The LRU providing the resolutionadvisory display function(normally coupled to the VSIdisplay on the VSI or EFISsystem) is not responding toTCAS inputs.

Verify wiring and power to the RA displays.Check that RA valid discretes 1 and 2 (RMP-14C and RMP-13E) are 10 ohms to groundmaximum. Repair wiring, or replace RA display,as required.

RAD ALT One or both of the radio altimeterinputs to TCAS have failed.

Verify wiring and power to the radio altimeter.Check that RAD ALT 1 and 2 are valid: RMP-2Kand RBP-3C should be at 28 V dc, orRMP-13H/J and RBP-3D/E should be validARINC 429. Repair wiring, or replace radioaltimeter, as required.

XPDRBUS

Both of the transponderinterfaces to TCAS have failed.

Verify wiring and power to the transponder.Check for data on TX 429 buses 1 and 2(RMP-14F/G and RMP-14H/J). Repair wiring, orreplace transponder(s), as required.

ATT Attitude input function is notused.

N/A



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Table 8.2. Antenna Wiring Resistance

ANTENNA CONNECTOR SECTION PIN DC RESISTANCE

Top Directional LTP 1

2

3

4

1000 ± 100 ohms

8060 ± 800 ohms

4020 ± 400 ohms

2000 ± 200 ohms

Bottom Omnidirectional LM?P 1

2

3

4

0 (50 ohms max)

Infinite (>50k ohms)



Bottom Directional LMP 1

2

3

4

1000 ± 100 ohms

8060 ± 800 ohms

4020 ± 400 ohms

2000 ± 200 ohms



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Table 8.3. CU Self-Test Execution

TEST SEQUENCE ACTIVATION TEST INDICATIONS

Power-on self-test Self activated witheach application ofsystem power.

a) All front-panel lamps come on during the three-second lamp test.

b) If the TCAS is operational, the TCAS PASSgreen lamp comes on for a 10-second displayperiod, and then goes off.

c) If the TCAS is not operational, one or morefault lamps come on for the 10-second displayperiod. Refer to table 8.1 for corrective action.

Continuous self-test

Executedautomatically as partof normal TCAS in-flight operation.

No indication, unless a fault is detected. Systemstatus/fault data is stored in memory for 10consecutive flight legs. Data may be recalled byperforming commanded self-test on the ground.

Commanded self-test

1) Press front panelPUSH TO TESTbutton.

Same as power-on test sequence

2) Press PUSH TOTEST button againbefore theprevious10-second displayperiod haselapsed.

a) Previous fault display is aborted.

b) All lamps come on during a one-second lamptest,

c) Status/fault data recorded during the precedingflight leg is displayed for 10 seconds.

3) Press PUSH TOTEST buttonbefore the end ofeach succeedingdisplay period.

a) Status/fault data recorded during a total of10 flight legs (maximum) is displayed.

b) (b)When test data for the earliest recordedflight leg has been displayed, all indicators willflash at a 2.5-Hz rate for three seconds ifPUSH TO TEST is again pressed. Thisindicates the end of recorded test data.



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9. EICAS CHECK

9.1 REVERSIONARY CONTROL

Step 1: With power applied to the system, observe that the EICAS display is displayingvalid data for both the No. 1 and No. 2 engines and systems.

Step 2: Open the circuit breaker for DAU 1A and observe that the data displayed on theEICAS for the left-hand systems becomes invalid.

Step 3: Press the DAU 1 pushbutton on the EICAS REV panel and observe that the EICASdisplay is now showing valid data for the No. 1 systems from DAU 1B.

Step 4: Close the circuit breaker for DAU 1A and open the circuit breaker for DAU 1B.Observe that the data displayed on the EICAS for left-hand systems becomesinvalid.

Step 5: Close the circuit breaker for DAU 1B and press the DAU 1 pushbutton on theEICAS REV panel to select DAU 1A for display and verify that valid data isdisplayed for the No. 1 Systems on the EICAS display.

Step 6: Open the circuit breaker for DAU 2A and observe that the data displayed on theEICAS for the right-hand systems becomes invalid.

Step 7: Press the DAU 2 pushbutton on the EICAS REV panel and observe that the EICASdisplay is now showing valid data for the No. 2 Systems from DAU 2B.

Step 8: Close the circuit breaker for DAU 2A and open the circuit breaker for DAU 2B.Observe that the data displayed on the EICAS for right-hand systems becomesinvalid.

Step 9: Close the circuit breaker for DAU 2B and press the DAU 2 pushbutton on theEICAS REV panel to select DAU 2A for display and verify that valid data isdisplayed for the No. 2 Systems on the EICAS display.



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9.2 CREW ALERTING SYSTEM

9.2.1 TONE GENERATOR AND MESSAGE ACKNOWLEDGMENT

Step 1: Simulate a Red Warning Message. Verify that the appropriate message isdisplayed as blinking red in the CAS message field on the EICAS display. Verifyalso that the warning tones are present from the aural warning system.

Step 2: Acknowledge the message by pressing the pilot’s Master Warning Pushbutton.Verify that the message goes steady (stops blinking) and that the warning tonestops.

Step 3: Simulate the Red Warning Message again. Verify that the appropriate messageis displayed as blinking red in the CAS message field on the EICAS display.Verify also that the warning tones are present from the aural warning system.

Step 4: Acknowledge the message by pressing the copilot’s Master WarningPushbutton. Verify that the message goes steady (stops blinking) and that thewarning tone stops.

Step 5: Simulate an Amber Caution Message. Verify that the appropriate message isdisplayed as blinking amber in the CAS message field on the EICAS display.Verify also that the caution tones are present from the aural warning system

Step 6: Acknowledge the message by pressing the pilot’s Master Caution Pushbutton.Verify that the message goes steady (stops blinking) and that the caution tonestops.

Step 7: Scroll the amber message off screen by using the knob on the EICAS bezel.

Step 8: Simulate another Amber Caution Message. Verify that the appropriate messageis displayed as blinking amber in the CAS message field on the EICAS display.Verify also that the caution tones are present from the aural warning system.Also verify that the first Amber message from step 5 above returned to the CASfield.

Step 9: Acknowledge the message by pressing the copilot’s Master Caution Pushbutton.Verify that the message goes steady (stops blinking) and that the caution tonestops.

Step 10: Simulate a Blue Advisory Message. Verify that the appropriate message isautomatically acknowledged after 5 seconds.

Step 11: Scroll the blue message off screen by using the knob on the EICAS bezel.

Step 12: Simulate another Blue Advisory Message by. Verify that the appropriatemessage is automatically acknowledged after 5 seconds. Also verify that thefirst Blue message from step 10 above returned to the CAS field.



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9.2.2 TAKE OFF INHIBIT FUNCTION

Step 1: Place the system in Takeoff Inhibit Mode, which is:

V = V1 < 15KTS and RA < 400 FT and Elapsed time of TO Inhibit < 60 SEC

Step 2: Simulate a Red or Amber message that is inhibited during Takeoff.

Step 3: Verify that the message is inhibited, i.e., not displayed in the CAS message fieldfor 60 seconds.

Step 4: Verify that after approximately 60 seconds the message is displayed in the CASfield and the aural warning is active.

9.2.3 RMU ENGINE BACKUP

Step 1: Press the PAGE button on the RMU.

Step 2: Select the line select key next to ENGINE.

Step 3: The RMU Backup Engine Page No. 1 should now be displayed.

Step 4: Press the Page 2 button.

Step 5: The RMU Backup Engine Page No. 2 should now be displayed.

Step 6: Press the PGE button on the RMU and then select the Return to Radios lineselect key.

Step 7: The Main Radio Page should now be displayed.

Step 8: Pull the C.B. for the left IC-600.

Step 9: RMU1 should automatically display the Engine Backup Page No. 1.

Step 10: Reset the EICAS display C.B.

9.2.4 RMU NAVIGATION BACKUP

Step 1: Press the PAGE button on the RMU.

Step 2: Select the line select key next to NAVIGATION.

Step 3: The RMU Backup Navigation Page should now be displayed.



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10. FLIGHT DIRECTOR TEST

Table 10.1. Flight Director Test

TITLE TEST NO. WORK STEPS

FDCMD

CONTROLWHEEL RESPONSE

Power Check 10.1.A. Check all plugs for voltages oncorrect pins

Use appropriate interconnect todetermine pins & voltages

Go-Around &WNDSHRModes

10.2.A. Push & release pilot’s Go Aroundbutton with WOW

Up TO annunciation on PFD will light.Command Bars will show 12 degnose up.

B. Engage AP with WOW. TO mode is reset.

C. Disengage AP.

D. Push & release copilot’s GoAround button with WOW.

Up TO annunciation on PFD shall light.Command Bars will show 12 degnose up.

E. Engage AP with WOW.

F. Simulate air mode, and greaterthan 400 ft radio altitude.

TO mode is reset.

G. Push & release pilot’s Go Aroundbutton with A/C in Air Mode.

Up

H. Engage AP. The internal monitorsmay disengage the AP shortlyafter engagement.

GA annunciation on PFD will light.GA mode remains engaged.

I. Push & release pilot’s Go Aroundbutton. Press AP DISC switch oncontrol wheel to reset AP.

J. Push & release copilot’s GoAround button with A/C in AirMode.

GA mode extinguishes

Engage AP. The internal monitorsmay disengage the AP shortlyafter engagement.

GA annunciation on PFD will light..

K. Disengage AP. Note that internalmonitors may disengage the APsoon after engagement. If thisoccurs, press the AP DISC switchon the control wheel to reset thesystem.

GA mode remains engaged.AP OFF is annunciated on EICAS.If monitors disengaged AP then APFAIL is annunciated until AP DISCis pressed.

L. Select GA with WOW.

M. Engage Test Mode forGPWS/Windshear with switchlocated on Maintenance Panel.Verify that windshear warning isannunciated on the PFD

TO annunciation on PFD will light.Command Bars will show 12 degnose up

N. Press GA again. WDSHR will be annunciated inplace of TO Mode.WDSHR is deselected.



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FDCMD


AnnunciatorLights Test

10.3.A Reference maintenance page1-35 of IC-600 self-test andperform lamp test.

HDG Select& FD OFFModes & CPLSelect

10.4.A. Connect pitot-static tester topilot’s pitot input. Set Hdg Bug onPFD to fore Lubber Line. SelectHDG on GC-550 FD modeselector.

Left HDG annunciators on GC-550 FDmode selector & PFD will light. FDpitch command in pitch sync mode.

B. Decrease Hdg Bug reading 10deg less than aircraft heading.

FD Command bar follows Hdg Bug.

C. Increase Hdg Bug reading to 10deg greater than aircraft heading.

Right FD Command bar follows Hdg Bug.

D. Increase Hdg Bug reading to 90deg greater than aircraft heading.

FD Command bar follows Hdg Bug.

E. Push & release BNK button onautopillot GC-550 controller.

Right Autopilot GC-550 controller LOWbank Limit Light turns on. PFDLOW Bank Limits are displayed.

F. Push and release BNK button onGC-550 autopilot controller,

Right Autopilot GC-550 controller LOWBANK Limit Light turns off. PFDLOW Bank Limits are removed.

G. Press Hdg SYNC to synch HdgBug reading with lubber line.

Null FD command bar follows Hdg Bug.

H. Increase altitude to 25,000 ft. Null At approximately 25,000 ft. theLOW BANK light on the autopilotGC-550 controller will light and theLow Bank will be annunciated onthe PFD.

I. Decrease altitude to 24,500 ft Null At approximately 24,750 ft. theLOW BANK on the autopilot GC-550controller and the Low Bankannunciation on the PFD willextinguish.

J. Set simulated airspeed to zero.Push and release FD Off buttonon the GC-550.

Null FD command bar is removed onappropriate PFD.

K. Reset HDG mode. HDG annunciators on GC-550 andPFD are removed.

L. Sel FD ON and push and SelectHDG on GC-550 FD modeselector.

HDG annunciators on GC-550 FDmode selector & PFD will light. FDcommand bars appear on PFD.

M. Push & release CPL button onGC-550.

CPL arrow on GC-550 and PFDschanges direction. HDG mode willextinguish.

N. Select CPL on GC-550 to returnCPL arrow to pilot side.

CPL arrow on GC-550 and PFDspoints left.



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FDCMD


NAV SelectMode, VORCourse

10.5.A. Simulate 100 kts airspeed. Tunepilot’s nav receiver to a VORfrequency. Set course Pointer tozero left-right (CDI) needle.Select NAV on FD MODESELECTOR GC-550.

Left Green NAV annun on FD MODESELECTOR GC-550 shall light.Green VOR Annunciator on PFDshall appear.

B. Decrease Course Pointer reading10 deg.

Right FD Command bar follows CoursePointer.

C. Increase Course Pointer reading10 deg from original position.

Left FD Command bar follows CoursePointer

D. Set simulated airspeed to zero.Reset all FD modes.

Null

NAV SelectMode, VORRadioDeviation

10.6.A. Simulate 100 kts airspeed. SetCourse Pointer 90 deg right ofaircraft heading. Simulategreater than 2 dots dev fly left, 25DME miles & a “TO” the stationindication. Set Hdg Bug on PFDto fore lubber line. Select NAVon FD mode selector GC-550.

Null HDG annunciators on FD modelselector GC-550 & PFD shall light.Green NAV annunciator on FDmode selector shall light. WhiteVOR annunciator on PFD shallappear. FD pitch command in pitchsync mode.

B. Reduce radio deviation slowly. Right NAV mode captures at approx. 1/2dot. HDG annunciators turn off.Green VOR annunciator on PFDshall appear.

C. Reset NAV mode. (Push NAVbutton on GC-550) Simulate 250kts airspeed. Simulate greaterthan 2 dots dev fly left.

Nav mode annunciators on PFD &mode selector GC-550 shall go out.FD Command bar is out of view.

D. Select NAV on FD mode selector Null HDG annunciators on FD modeselector GC-550 & PFD shall light.NAV annunciator on FD modeGC-550 selector shall light. WhiteVOR annunciator on EADI PFDshall appear. FD pitch command inpitch sync mode.

.E. Reduce radio deviation slowly. Right NAV mode captures at about 1 dot.HDG annunciators turn off. GreenVOR annunciator on PFD shallappear.

F. Reset all FD modes. FD Command bar is out of view.



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FDCMD


NAV SelectMode, VORAPP & OSS

10.7A. Simulate 100 kts airspeed. SetHdg Bug & Course Pointer onPFD to fore lubber line. Simulategreater than 2 dots dev fly right,10 DME miles & a “TO” thestation indication. Select APR onFD mode selector GC-550.

Null CTR HDG annunciator on FD modeselector GC-550 & EADI PFD shalllight. White VAPP annunciator onPFD shall appear. FD pitchcommand in pitch sync mode.

B. Reduce radio deviation to zero.Wait 20 seconds before startingtest no. 7C.

Nulled HDG annunciators turn off. NAVannunciator on FD mode selectorGC-550 shall light. Green VAPPannunciator on PFD shall appear.Allows OSS circuit to time out.

C. Pull C.B. for AZ-850.

D. Rapidly rotate course pointer plus& minus 30 deg from aircraftheading.

OSS circuit trips. FD command barfollows course pointer change, notradio dev.

E. Set course pointer 5 deg greaterthan aircraft heading & 1 dotradio dev fly left

Right For about 30 seconds FD commandresponds only to course error thenmoves left in response to radio devsignal. Checks OSS circuit timeout.

F. Reset APR on FD mode selectorGC-550. Set course pointer 10deg less than aircraft heading &radio dev to zero.

Nav mode annunciators on PFD &mode selector GC-550 shall go out.FD Command bar is out of view.

G. Reset all FD modes. FD Command bar is out of view.

NAV SelectMode,Localizer

10.8.A. Tune pilot’s nav receiver to aLOC frequency. Set CoursePointer 90 deg right of aircraftheading. Simulate greater than 2dots dev fly left. Set Hdg Bug onPFD to fore lubber line. SelectNAV on FD mode selectorGC-550.

HDG annunciators on FD modeselector GC-550 & PFD shall light.NAV annunciator on FD modeselector GC-550 shall light. WhiteLOC annunciator on PFD shallappear. FD pitch command in pitchsync mode.

B. Reduce radio deviation to zero. Right HDG annunciators turn off. GreenLOC annunciator on PFD shallappear.

C. Set Course Pointer to aircraftheading.

Left

D. One minute after test no. 8.C.rotate Course Pointer 20 deg leftof aircraft heading.

Left FD command initially goes left thennulls. Checks phasing & crosswindwashout for LOC.

E. Rotate Course Pointer 20 degright of heading.

Right FD command initially goes rightthen nulls. Checks phasing &crosswind washout for LOC.

F. Return Course Pointer to aircraftheading.

FD Command bar is out of view.

G. Reset all FD Modes



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FDCMD


APR SelectMode, GlideSlope

10.9.A. Tune pilot’s nav receiver to aLOC frequency. Simulate 150 ktsairspeed. Simulate 1500 ft onradio altimeter by putting piece ofEcho Sorbs on either RadioAltimeter Antennas. Set Hdg Bug& Course Pointer on PFD to forelubber line. Simulate zero LOCdeviation. Simulate 1 dot fly upGS deviation. Select APR & ALTon FD mode selector GC-500.

NAV, APR & Alt annunciators on FDmode selector GC-550 shall light.Green LOC, ALT & white GSannunciators on PFD shall appear.AP & YD ENGAGE lights turn on.

B. Reduce GS deviation to 1/4 dotfly up.

DnthenUp

Green GS annunciator on PFD shalllight. Initial command is fly down.After 15 seconds command is flyup.

C. Reduce GS deviation to zero. Nulls

D. Simulate 1/4 dot fly down GSdeviation.

Dn FD command goes down.

E. Open radio altimeter valid or pullradio altimeter circuit breaker.Reestablish original conditions intest no. 9.A.

NAV CAP, APR ARM & ALTannunciators on FD mode selectorGC-550 shall light. Green LOC,ALT & white GS annunciators onPFD shall appear.

F. Reduce GS deviation to zero thenreturn it to 1/4 dot fly up.

DnthenUp

Green GS annunciator on PFD shallappear. Initial command is flydown. After 15 seconds commandis fly up.

G. Reset ALL FD modes. Setsimulated airspeed to zero

FD Command bar is out of view

BC SelectMode, BackCourse

10.10.A Tune pilot’s nav receiver to aLOC frequency. Set CoursePointer on PFD to aft lubber line(tail of course pointer is forward).Simulate greater than 2 dots devfly left. Set Hdg Bug on PFD tofore lubber line. Select APR &ALT on FD mode selectorGC-550.

Null HDG, BC & ALT annunciators onFD mode selector GC-550 shalllight. Green HDG, ALT & white BCannunciator on PFD shall appear.

B. Reduce radio deviation to zero. HDG annunciators turn off. GreenBC annunciator on PFD shallappear.

C. Rotate tail of Course Pointer 10deg right of aircraft heading.

Right FD command initially goes rightthen nulls. Checks phasing &crosswind washout for BC.

D. Return tail of Course Pointer toaircraft heading.

E. Reset all FD modes. FD Command bar is out of view.



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FDCMD


ALT SelectMode,Altitude Hold

10.11.A. Set Hdg Bug on PFD to forelubber line. Select HDG & ALTon FD mode selector GC-550.

Null HDG & ALT annunciators on FDmode selector GC-550 & EADI PFDshall light.

B. Simulate 200 ft greater thanpresent altitude.

Down Checks phasing of ALT error signal.

C. Deselect & reselt ALT onGC-550.

Null System resynchronizes to newaltitude at ALT engage.

D. Simulate 200 ft less than originalaltitude.

Up Checks phasing of ALT error signal.

E. Reset all FD modes. DisengageAP.


ALT SELMode AltitudePreselect &DepartureWarning

10.12.A. Preselect 10,000 ft using ALTSELknob. Simulate 2000 fpm climbwith ADC tester. Set Hdg bug onPFD to fore lubber line. SelectHDG and VS on FD modeselector GC-550.

HDG & VS annunciators on FDmode selector GC-550 light. GreenHDG, VS and white ASELannunciators appear on PFD. 2000fpm air data command appears onPFD.

B. Allow altitude to increase towards10,000 ft.

Null At approximately 9000 ft the ALTPRESELECT WINDOW on the PFDchanges from blue to amber, andthe altitude alert horn sounds for 1second.

Down At approximately 9600 ft the VSannunciator on the FD modeselector GC-550 turns off and theASEL arm annunciator on the PFDchanges from white to green. Awhite box around the green ASELannunciator is displayed for 7seconds indicating mode capture.

At approximately 9750 ft the amberaltitude select window changesfrom amber to blue.

C. Reduce Climb Rate to zero at10,000 ft.

Null The green ASEL annunciator onPFD turns off and the green ALTannunciator on the FD modeselector GC-550 and the PFDlights.NOTE: System will not go into ALThold mode if rate of climb is notless than 300 FPM and within 25 ftof selected altitude.

D. While in ALT HLD decreasealtitude below 9750 ft

Up At approximately 9750 ft theAltitude select window changesfrom blue to amber and a hornsounds for one second for thealtitude departure warning function.



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FDCMD


VS SelectMode,VerticalSpeed Hold

10.13.A. Simulate 2000 fpm climb withPitot/ Static tester. Set Hdg Bugon EHSI to fore lubber line.Select HDG & VS on FD modeselector GC-550.

Null HDG & VS annunciators on FDmode selector GC-550 & PFD turnon. 2000 fpm air data commandappears on PFD.

B. Increase VS climb rate to 2500fpm.

Down Checks phasing of VS error signal.

C. Push & hold TCS button untilsignal nulls

Nulls Synchronizes to new VS signal. Airdata command displays dasheswhile TCS button pushed, new VSwhen button released.

D. Decrease VS climb rate to 1000fpm.

Up Checks phasing of VS error signal.

E. Push & hold TCS button untilsignal nulls.

Nulls Synchronizes to new VS signal. Airdata command displays dasheswhile TCS button pushed, new VSwhen button released.

F. Move pitch wheel on APcontroller fore and aft

Resets VS hold mode. VSannunciators on FD mode selectorGC-550 & PFD turn off. FDcommand in pitch sync mode.

G. Reset all FD modes. Setsimulated VS rate to zero.




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FDCMD


SPD Mode 10.14.A. Simulate 15,000 ft altitude & 200kts airspeed with ADC tester. SetHdg Bug on PFD to fore lubberline. Select HDG & SPD on FDmode selector GC-550.

Null HDG & IAS annunciators on FDmode selector GC-550 turn on.HDG & IAS annunciators on PFDappear. 200 kts air data commandappears on PFD.

B. Decrease airspeed to 190 kts. Down Checks phasing of IAS error signal.

C. Increase airspeed to 210 kts. Up Checks phasing of IAS error signal.

D. Push & hold TCS button untilsignal nulls.

Nulls Synchronizes to new IAS signal.Air data command displays dasheswhile TCS button pushed, new IASwhen button released.

E. Move pitch wheel on APcontroller in climb or descend.

Sync Resets IAS hold mode. IASannunciators on FD mode selectorGC-550 & EADI PFD turn off. FDcommand in pitch sync mode.

F. Select SPD mode on GC-550. Null IAS annunciator on FD modeselector GC-550 turns on. IASannunciator on PFD appears. 200kts air data command appears onPFD.

G. Increase altitude slowly to 26,000Ft.

As altitude passes thru 25,100,MACH is annunciated on theGC-550 and the PFD. The air datacommand displays a Mach numbertarget.

H. Decrease altitude slowly to24,000 ft

As altitude passes thru 24,900 ft,IAS is annunciated on the GC-550and the PFD. The air datacommand displays an airspeedtarget.

I. Reset all FD modes. Setsimulated altitude to ambient &airspeed to zero.




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11. AUTOPILOT TEST

Table 11.1. Autopilot Test


FDCMD


NOTE: Use appropriate interconnect to detremine pins & voltages

AP and YDEngageDisengageLogic

11.1.A. Push AP Engage button onGC-550 Guidance Controller

AP and YD Annunciators on GC-550are ON and AP and YD indicationsare on the PFD..

B. Push AP Engage button onGC-550 Guidance Controlleragain

AP button annunciator turns OFF onthe GC-550 and the YD annunciatorremains ON. The AP indication ofthe PFD flashes for 5 sec and turnsOFF. The YD indication remainsON.

C. Push the YD button on GC-550Guidance Controller

The YD annunciator on the GC-550turns OFF and the YD indication onthe PFD turns OFF.

NOTE: Engage the AP and YD with the AP button on the GC-550 before each of the following work steps D thru J.

D. Activate pilots AP/TRIM/PUSHERDISC switch

AP and YD annunciators on theCG-550 are ON AP and YDindications are on the PFD.

E. Activate copilotsAP/TRIM/PUSHER DISC switch

Aural warning system sounds. APand YD flash for 5 seconds on thePFD, then extinguish. AP and YDannunciators on the GC-550 turnOFF.

F. Activate the pilots UP/Down TrimSwitch

Aural warning system sounds. APflashes for 3 seconds on the PFD,then extinguish. YD remains ON.AP annunciators on the GC-550turn OFF and YD remains ON.

G. Activate the copilots UP/DownTrim Switch


H. Activate the backup test switchon the pedestal in UP/Down orDown position


I. Pull the IC 1 Circuit Breaker Aural warning system will sound.On the copilots PFD, AP and YDflash for 5 seconds then extinguish.FD failure will be ON. YD remainsON. The pilots PFD, MFD, andEICAS will show Red X..

J. Pull AP 1 Circuit Breaker Aural warning system will sound.On the PFD AP (red) and YD(amber) will flash for 5 YDannunciators are OFF. seconds andthen remain ON until the failure iscleared. ON the GC-550 AP and



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FDCMD


Refer to Note 1 in the introduction section prior to performing the following tests.

Autopilot RollAxis

11.2.A Set Hdg Bug on PFD to forelubber line. Select HDG on FDmode selector GC-550. Centercontrol wheel & engage AP.

HDG annunciators on FD modeselector GC-550 and PFD shalllight. AP & YD ENGAGE lights turnon.

B. Rotate TURN knob on autopilotcontroller full clockwise.

CW HDG annunciators turn off.Autopilot responds to turn knobcommand.

C. Rotate TURN knob fullcounterclockwise.

CCW Autopilot responds to turn knobcommand.

D. Disengage autopilot with APTRIM DISC switch.

AP & YD ENGAGE lights turn off.

E. Push & release AP ENGAGEbutton.

AP & YD ENGAGE lights remainoff.

F. Rotate TURN knob to centerdetent position. Engage AP.

CTR AP & YD ENGAGE lights turn on.

G. Depress and hold A/P TCSbutton. Turn control wheel rightand then left and then centercontrol wheel.

CTR AP & YD remains engaged.

H. Release A/P TCS button.

I. Disengage AP & YD. AP & YD ENGAGE lights turn off.

AutopilotPitch Axis

11.3.A Select ALT on FD mode selectorGC-550. Center control column &engage AP.

CTR ALT annunciator on GV-550 FDmode selector & PFD shall light.AP & YD ENGAGE annunciate onPFD.

B. Rotate PITCH wheel on autopilotcontroller towards DESCEND.

FWD ALT annunciators turn off.Autopilot responds to pitch wheelcommand.

C. Rotate PITCH wheel on autopilotcontroller towards CLIMB.

AFT Autopilot responds to pitch wheelcommand.

D. Depress A/P TCS and recentercontrol column. Release A/PTCS.

CTR Control column free with TCSdepressed. Upon releasing TCSautopilot in pitch old mode.

E. Disengage AP. AP & YD ENGAGE lights turn off.



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FDCMD


AutopilotTrim

11.4.A Using Aircraft UP/Down TrimSwitches located on Yoke, runtrim to NOSE UP limit. Movecontrol column slightly forward &engage AP.

AP & YD ENGAGE lights turn on

B. Rotate PITCH wheel on autopilotcontroller 2 turns towardsDESCEND.

FWD After 3 to 4 seconds autopilot trimdrives nose down. Approximately 5seconds later Pitch Trim Indicatoron EICAS will move is a pitch downdirection. .Verify trim is running in the properdirection.

C. Disengage AP. AP & YD ENGAGE lights turn off.

D. Using Aircraft UP/Down TrimSwitches located on Yoke, runtrim to NOSE DOWN limit. Movecontrol column slightly aft &engage AP. Rotate Pitch Wheelon Autopilot Controller 2 turnstoward Climb.

After 3 to 4 seconds autopilot trimdrives nose down. Approximately 5seconds later Pitch Trim Indicatoron EICAS will move is a pitch downdirection. .Verify trim is running in the properdirection.

E. Disengage AP & reduce airspeedto zero

AP & YD ENGAGE lights turn off.

avionic emb 145-3

Documents

fore lubber

teps fd cmd

fd command

fd command

reduce gs

pitch wheel

line select

customer maintenance